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AIRP~"'E

DESIGN. TENI'H EDITION. 1954

CONTENTS

Tote1 Pages

Chapter

1 2

3 L~

5

6 ~,

{

8

9

10 11

Preface............................................. Foreword "Airplane Design Made Simple".............. Notation and Abbreviations..........................

1

Introduction..... ..•..•...•.••••..•.••....••......••

4

Layout Design of Light Airplanes •••••••••••••••••••• Layout Design of Transport and Cargo Airplanes •••••• Layout Design of Flying Boats ••••••••••••••••••••••• Layout Design of High Speed Airplanes ••••••••••••••• Layout Design of Helicopters •••••••••••••••••••••••• Structural Dosign Considerations •••••••••••••••••••• Production Methods and Costs as Factors in Design ••• Wing Design ••••••••••••••••••••••••••••••••••••••••• Control Surface Design ••••••••••••••••••••••••••••••

12

2

2

4

8 7 22

.3 2 15 16

Landing Gear Design ••••••••..•••••••••••••••..••••••

9 17

Fuselage and Hull Design ••••••••••••••••••••••••••••

10

Total Text Pages ••

161

Design Data Appendices Appendix 1

Properties of Air and Airflow Data. Including Super-

2

Aerodynamic Data a. Wing Characteristics and Dimensions ••••••••••• b. Fuselage Drag Data •••••••••••••••••••••••••••• c. Power Plant Drag Data ••••••••••••••••••••••••• d. Control Surface Design Data--Stability and Control ••••••••••••••••••••••••••••••••••.•• e. Propeller Data •••••••••••••••••••••....•••• , .. f. Ferformance Charts and Data •••••••••••••••• u • •

sonic Flow Data...................................

3 4

Hydrodynamic Data ••••••••••••••••••••••••••••••••••• Power Plant Data •••••••••••••••••••••••••••••••••••• iTeight Data ••••••••••••.•.•• • ,••••••••••.••••.•••••••

6

structural Design Data

5

s. b.

7 8

.Applied Loads •••••••••••••••.•••••••••••.••••• Strength Data ••••••.•••••••••••••••••.••••••••

Cost Data ••••••••••••••••••••••••••••••••.•••••••••• Data for Pass~nger Accomodations and Comfort ••••••••

13

34

14 6 6

14 14 7 10

38

28

29 9 4

Answers to Problems............

3

InC. ex • • • • • • • • • • • • •• • • • • • • •• • • • •

6

Errata.........................

1

Pac~s •••••••••••

236

Total Pages •.••••••••••••••••••

397

Total Appendix

P:l PREFACE TO TEIITE EDITION This book aims to show how an airplane or helicopter can be designed and to include in appendix ferm most of the data necessery for a student to carry a design project through it s preliminary stages. The book aims to provide occasion for applying fundBIDentals of fluid mechanics, thermodynamics, aerodynamics, strength of materials, and airplane structures to a design problem of current and future practical importance. With this aim frequent revisions of the text are imperative, expecially since the fundamental units going to make up an airplane are in a state of rapid development. 'tiith frequent revision needed, photo-lithographing, which is economically feasible in lots of under 2,000 copies, is the only reasonable method of presentation. This book was revised nearly every year from 1934 to 1943, when t he seventh edition ",as released. Majer revisions prior to 1943 were fortunately unnecessary because the author concentrated on probable future trends rather than current practice and made a fortunate early appraisal of the importance of the retractable tricycle landing gear long before its widespread adoption. During the war years, from 1943 to 1947, no new editions could profitably be issued because all important developments leading to a reasonable appraisal of the components of the airplane of the future were shrouded in military secrecy. In the Bpring of 1947, thousands of formerly classified documents were released and it again became possible to discuss public ly, with substant iating technical data, the trends in airplane design which may reasonably be expected in t he near future. The eighth edition was issued at that time. Civil Aeronautic Regulations also underwent a major change in 1947 thru the promulgation of CAR 3 modifying the older CAR 04. The ninth edition in 1949 made further use of post-war information releases. As this tenth edition goes to press the rate of release of new technical data is at an all-time high and the need for a coordinated presentation greater than ever, but the field has grown so diverse as to be confusing in its complexity and e well indexed NACA file is more than ever indispensible to any adequate preliminary design study. Many missile developments are of course still security classified, but sufficient basic information is now available for useful predesign studies by students and other amateurs. The studant, even more than the industrial preliminary design engineer, must design for the future, as his expected industrial work is conSiderably farther in the future than that of the industrial engineer. Recent developments in jet propulsion, lOi'; drag, and high lift permit a vision of future commercial aircraft racically different from any available on the market today. Such expected developments must be considered in this text in order for it to be useful. This text is complementary to the textbook, Technical Aerodynamics, second edition (r.:cGraw-Hil1. 1947, also distributed by the University Bookstore) by the same author and is intended for use ~Jith the follo\.;ing U.S. Government publications: (a) Strength of Metal Aircraft Elements, Publication ANC-5a (Government Printing Office, $1.25), (b) Airplane Alrworthiness, Publication CAR-3 which is distributed by the U. S. Government Printing Office. Those parts of CAR-3 which relate to airplane design appear in this text for the convenience of the student as pages A6a:l to pA6a:26 inclusive. Numerous tables from ANC-5a (reproduced with permission of the Munitions Board, Aircraft Committee) are elso included in the appendix material (A~pendix A6b). Particular the.nks ure cue to the National Advisory Comrd ttee for Aeronautics and to McGrew-Eill publications for permission to reproduce figures and data, and to Mr. and Mrs. Jack Burnell and Iv~rs. Kathleen Zylka for assistance in preparation of the tenth edition. Suggestions for revision would be greatly appreciated by the author. K •. D. Wood Boulder, Color., A.pril 1954

p:2

FOREl~RD

The study of airplane desi~n must inc~ntally be a study of airplane designers, and should preferably include a portrayal of the conditions under which commercial airplanes are commonly developed. The following articles from Aviation by R. R. Osborn written prior to

1940, while no longer as applicable as at the time of their original writing present some personnel aspects of the development engineering problem which are still, with important variations, involved in many development projects.

AIRPLANE DESIGN MADE SIlIPLE

"Lately we have been very much surprised to find that airplane deSign and construction seem to be very mysterious to some people associated directly with the industry, as well as to the general public. 'r.ley have no idea why a biplane is used for one type of airplane and a monoplane for the next type. 7hey probably wonder why the enPine installed was selected, and why the cabin or cockpits are arran~ed as they are. In fact, in some cases they have even wondered why the airplane was ever built. RealiZing that some information along this line would probably be appreciated by our readers, we have interviewed a number of experienced desi~ers we know, to learn from them the reasoning and processes by means of which a new airplane is created. They were glad to tell us their experiences and we have condensed all of their stories into the following, which might be said to be the high points in the life of an average airplane in its journey from the drafting board to the field: "As his favorite layout draftsman is working up some advertiSing for the sales department, the Desi~ner is much discourared to find that he will have to use an inexperienced man and do the figurinp' and calculating himself. "Desi~er calls for a win" span of 37.5 feet. layout draftsman misunderstands his writing and lays out the airplane to have 375 sq. ft. of wing area. "Airplane originally laid out as a monoplane. Ne'll" Department of Commerce Inspector shifted to the district. New Inspector has a great preference for biplanes, so design is changed to a biplane. "President sends in word that speed is essential in all ne'll" aircraft of the immediate future, and airplanes must be desi~ed mainly for speed. Design is altered to suit. "Engine selected is the one manufactured by the Chief En"ineerts Golfing partner. Designer asks the world howinell he can turn out a good Ship when he has to use an en~ine like that one? Chief Engineer's gol! game gets poorer so that his partner beats him regularly. DesiVler ordered to shift to the best en!7ine available in another company. Desi!7fier asks the 'll"orld howinell he can turn out a /"ood ship around an engine like that one? "President sends in a note stating that the _tcllword is economy, and that all ne'll" designe should have cheapness of construction and economy of ~eration as their major criteria. Desi~ is altered to suit. "DeSigner hears that the Whoosis Airplane Company is la:rin~ out a competin!7 model with gull-shaped wings. I:!mediately scraps his design and starts over again 1I'ith gull-whaped 1I'in~s. Simultaneously, the desi"ner of the \1hoosis Airplane Company has scrapped his drawinp,s and starts ne'll" layouts using butterny- shaped wings, after hearing that the Whatsis Airplane company is proceeding on that basis. "President returns from a tour arourd the country. Circulates notes to the effect that the present trend is toward better vision for the pilot, and that all other features, including speed and cheapness of construction, should be canpramised to obtain better vision for the pilot. Desi~ is altered to suit. "President sends in 'll"ord that the crying need of this country today is a good 5-cent oigar. Design is altered to suit.

"Shop makes an error, in buildin~ the fuselage a foot too short. In exchan~e for previous shop favor in covering up one of his errors, the desi~ner writes a long treatise to the Chief Engineer pointing out the trend to shorter fuselage lengths, suggesting that the fuselage be made shorter by 1 ft. Chief Engineer does not grasp the full meaning of the obscure part of the Designer'S calculations, so issues order to have the nose of the fuselage shortened by 1 ft. Designer and Shop Superintendent talk it over, and decide they had better just cut 1 ft. off of the nose and say nothing more about it. "Engine finally arrives for installation in the ship. Turns out that the engine company had decided to build a nine-cylinder engine instead of a sevencylinder engine. Long correspondenoe between airplane company and engine company to determine i f t'll"O cylinders shall be taken off or i f engine mount shall be chan?ed. Matter finally settled by flipping the coin. Engine mount is changed. "On installation of the engine it is found that the carburetor interferes 1I'i th the center landing gear fitting. Engine sent back to the engine plant to be made into a down-draft carburetor. When the engine returns it is discovered that the ne .... carburetor interfers with the oil tank. Send engine back to engine plant to be made over into a solid-fuel injection engine. "None of the shop c01l'l workers understanding English, Project Engineer _ves his arm around in the air to show them what type of wiD!! fillets he wishes. Thinking he is referring to the engine compartment C01l'l, they turn out a startlin!7 new idea in engine c01l'l. Project Engineer has dra1l'ing made to suit and sends dra1l'ing in to Chief Engineer pointing out that his new design will probably add 4 m.p.h. "Landin" ~ear was laid out for large diameter wheels. Somebody invents small diameter wheels and sells them to the PurchaSing Agent. When they are applied to the ship it is found that the propeller ground clearance is too small. Project Engineer announces that a three-blade propeller will be used because of high propeller tip speeds or something. "During set-up opetation, upper wing is found to interfere 1I'ith a beam in the roo! of the factory. After comparing costs of altering the beam in the roof or changing one set of 1I'ing struts, gap bet....een the Wings is decreased by 6 in. "First 1I'8if'hing of the ship ah01l's the center of gravity to be badly out of position. Upper 1I'ing is taken off and changed to one of large S1Ieep-back to balance the ship. Chief Engineer sends note to President explaining delay as necessary, as sweepback has to be used to improve pilot's vision. "At the field 1 ft. of left 1I'ing tip is knocked off on a hanger door. One foot is sawed off the other tip to match, and both ends are faired off neatly. "The airplane is put over the speed cOllrse and is found to have a high speed 5 m.p.h. more than the reSigner expected, but 5 m.p.h. less than he wrote in the preliminary specification. 'r.lis speed is 10 mop.he more than the Design Engineer expected and 10 m.p.h. less than he promised the President. The speed is 15 m.p.h. more than the Sales llanager expected and 15 m.p.he less than he wrote into the preliminary advertiSing copy. "Knowing his organization thoroughly, the speed is exactly what the President> antiCipated."

FOREWORD

P:3

AN AIRPLANE DESIGNER BEGmS A NEW PROJECT

"Having finished the morning paper the DeSigner leans back in his chair and starts to read over the customer's specification for the new airplane. "Thinks it would be a good idea to underscore with red pencil the parts of the customer's specification which will affect the design. After completing four pages finds that he has underscored all but three words so throws down specification in disgust. "Goes into Drafting Roan to discuss latest sporting news with favorite layout draftsman. Finds him busy on a rush job for ano~~er designer. Dashes into Chief Engineer's office and pounds on desk, demanding that favorite draftsman be transferred to his project and moved into his office to aSSist, as no other draftsman is able to understand what he wants done. Chief Engineer IZrunts and says that he'll think about it. "wanders through drafting room looking at work being done for other designers and offering suggestions which involve scrapping all drawinf,s and starting over again. "Designer is startled on returning to his office to find that favorite draftsman has already been moved in and is ready to go to work. "Suggests that centerlines be drawn here, here, and here, and returns to desk for contemplation. "Reads through specification hurriedly and then slams it down on ~esk asking howinell customer expects to get all that in one airplane. "Looks at drafting board and suggests that center lines be moved to here, here, and here to allow more roam for expansion of sketches. "Lights cigarette and starts reading specification again with determination. Discovers that latest ~ model of engine is called for. Swears blue streak but is secretly glad as draftsman will be kept busy for a few hours making a scaled-down drawing of engine. "Gets new notebook and paper filler from stock roan and letters name of new project and his name carefully on front cover, inking in letters with beautiful shading. "Places feet on desk and starts trying to concentrate on the details of the specification again. "Factory Superintendent calls up and says would like him to look at a fitting of his design which is giving trouble in shop. De signer says that he'll be down immediately to look at it. Shop Superintendent faints at other end of phone as he expected that Designer would manage to get down to see fitting in about three days, as usual. "Returns to offices and starts in on specifica-

tion again. Notices grasshopper on window sill. Studies unique details of grasshopper and considers application of catapulting gear for Navy ships. "Goes over to golf club for lunch and discusses merits of new design of clubs with profeSSional. "Returns to plant and as he passes watchman's gate-house hears important baseball game being broadcast on radio. Listens to several innings, discussing probable outcome of pennant race with watchman. "Back in office starts reading over specifications again. "Admires lettering on cover of new notebook and then numbers pages therein, using ornamental figures. "Suddenly realizes that i f he is to turn out design which 18 absolutely up-to-date it will be necessary for him to read up on latest developments here and abroad as noted in aeronautical magazines. Gets magazines and reads all social and political news therein. lIakes mental note to read technical articles later. "Wanders down into shop to watch operation of new rivetting machine. "Talks over international political situation wi th foreman of the Sheet Metal Shop. "Hears report that new airplane built by c~ peting company has landed at field so drives over to see i f there are any new ideas thereon to be appropriated. Looks ship over carefully. Points out to foreman of Hanger Cr_ all details which were improperly designed and expresses amazement that competitor managed to get a large production order on such a poor airplane. "Walks down to the School Hangar to watch students practicing landing. Comes to conclusion that modern landing gears are pretty good after all. "Back at office starts to read over specification again but notices that his slide rule is in need of cleaning. Decides hi' had better clean rule thoroughly as he will be using ita lot. "Also notices that desk drawer in which he keeps cigarettes, rubber bands, chewing gum, paper clips, smoking tobacco and pipe cleaners is in need of fixing up. Takes considerable care in working out good arrangement of contents. "Sees that it is almost quitting time and i f he doesn't hurry he will probably hold up the starting t:line of his golfing foursome. puts on hat and coat and l!a!lders over i' or look at drafting board. Observes that favorite draftsman has made progress on preliminary sketches for new design.

NOTATION AND ABBREVIATIONS

p:4

The following list of symbols used in this book is, with few exceptions, consistent with the practices in the United states of the N.A.C.A. and the Army-Navy-Comn~rce Committee on Aircraft Requirements. a a a a..c. A

AD b

B

br Bhp c

e.g. c.p.

C.

acceleration, ft./sec. 2 • slope of OTaph of Cr, vs. c;r (dCr/d cY ) per degree • position of aerodynamic center, fraction of chord; also subscript "actual" • aerodynamic center • area of cross section, sq. in. aspect ratio • equivalent drap area sq. ft. Also designated 2

h

by f.

• span of winff, ft.; also distance between spars, fraction of chord; also web thickness for spars, inches; width of sections; subscript "bendinf'''. • buoyant force, lbs; slenderness ratio factor (See Sq. 1:24) • subscript "bearing". a brake horsepower • subscript "chord"; coefficient; constant; generally fixity coefficient for columns; subsCript "compression". • center of ~avity • center of pressure, distance from leading edge. ft. • chord, ft.; cross wind force, lbs.; coeffiCient; constant; circumference.

• coefficients of lift, drag, and cross wind force (~ • L/qS (CD. D/qS (Cc • C/qS .. ideal minimum drag coefficient • induced drag coefficient CDo • profile drag coefficient • moment coeffiCient, about quarter chord unless CM otherwise specified • moment coeffiCient at zero lift Cyo • rate of climb, ft./min. Ch ). center of pressure, fraction of chord from leadCp C.P. ) ing edpe Cs • speed power coefficient for propellers d • diameter, ft.; also drag loadinp, lbs./sq.ft.; d • W/AD; depth or heieht; mathematical operator denoting differential. D s drag; Ibs.; diameter ft. e • ratio of wing weight to ,,-ross weil1'ht; unit deformation or strain; eccentricity; subscript for Euler's formula; Bubscript "endurance". E • efficiency; also chord ratio for tail surfaces; also modulus of elasiticity in tension .. modulus of elasticity in compression • flat plate area of CD • 1.00 equivalent to minimum parasite and profile drag of airplane. Also deSignated by An f .. unit stress, Ibs./sq.in.; also front spar location, fraction of chord also subscript "fusela"e"; internal -Cor calculated) stress • force, Ibs; allowable stress internal (or calculated) primary bending stress • internal (or calculated) precise bending stress • allowable bending stress, modulus of failure in bending • endurance limit in bending • internal (or calculated) bearing stress • ultimate bearing stress .. internal (or calculated) compressive stress • allo~ble compreSSive stress • ultimate compressive stress • compressive yield stress • proportional limit in compression • column yield stress • internal (or calculated) normal stress • allowable normal stress • internal (or calculated) shearing stress • allowable shearing stress • Factor of Safety • proportional limit in shear • modulus of failure in torsion • endurance limit in torsion

hp H

Hs HP i I

Ip j

~

m

• allowable stress in pure shear internal (or calculated) tensile stress • allowable tensiles stress • ultimate tensile stress • tensile yield stress • proportional limit in tension • acceleration due to gravity. 32.2 ft./ sec.• 2 • modulus of elasticity in shear, (modulus of rigidity) • altitude, ft.; also distance measured perpendicular to MAC as a fraction of MAC; heipht or depth; especially the distance between centroids of chords of beams and trusses. • horsepower • absolute ceilin~, ft. • service ceili~, ft. • rated horsepower • subscript "induced"; slope (due to bending) of neutral plane of a beam, in radians. (1 radian. 57.3 degrees). • moment of inertia of mass, slug-ft. 2 ; also moment of inertia of area, in.4 • polar moment of inertia • position of wing c.g., fraction of chord; also ~ , ft.; also !if; subsoript for Johnson'lI formula. -( ~ • torsion constant. • factor of safety; radius of gyration • span factor • coeffiCient, constant, or general factor • bearing factor of safety • lift, Ibs.; length, ft.; subscript "lift" or "level"; subscript "lateral". • parasite loadin a , lbs./sq.ft •• Wife Also deSignated by d • W/An • span loading, Ibs./sq.ft •• W/e(kIb)2 • thrust horsepower loading • w/Tho.. • mass, slugs; slope of lift curve-rdCt/dO() per radian; also subscript "maximum vertical", or "maximum".

n

• • • •

N

•

o

•

p

•

p

•

psi r R

• • • • •

p S

•

t

•

s

•

T

• • • • • • • •

!lIph

M

MAC

q Q

u U

•

•

• 11'

*

•

miles per hour moment, ft. Ibs.; also subscript "moment" mean aerod)7UUnic chord applied load in terms of W; rate of rotation, revs./sec.; subscript "normal". rate of rotation, revs./min.; also subscript "normal force" subscript "zero lift", "initial", or "standard". subscript "polar"; subscript "proportional l:iJni t"; power loading, Ibs./hp. • vr/p load (total, not unit load) J engine horsepower, design unless specified. pounds per square inch. dynamic pressure, Ibs./sq.ft •• 1/2/ ()v2 static moment of a cross section. radiUS; near spar location, fraction of chord resultant force or reaction, lbs.; subscript "resultant"J, stress ratio. Reynolds number. Vc ~~ surface (wing, unless otherwise noted; shear force. thickness; subscript "tail" or "terminal" or "tenSile". subscript "shear", "stalling"; wing loading, lbs.!sq.ft •• W/s thrust, lbs.; torsional moment, torque subscript "ultimate". gust velocity, ft./sec • airplane velOCity, miles per hour. airplane velocity, ft./sec. limited diving velOCity, ft./sec. design flap speed stalling velOcity, ft./sec. maximum speed of level flight, ft./sec. maximum theoretical diving velocity with zero propeller thrust, ft./sec. specific weight; unit pressure, lbs./sq.ft.; also subscript ''1r1ngl' average unit pressure, Ibs./sq.ft.; also gross weight.

in CAll 3 V dea1gna tea spe ad in lIIph.

§

NOTATION AND AEBREVIATIONs

x

• distance alon(' elastic curve of a beam; also

distance measured parallel to MAC in terms of ).(AC. x,y,z • axes, see Fig. PI1. y • denection (due to benc!1.rli'.) of elastic curve of a beam; distance from neutral axis to outer fibre; subscript "yield". Z • section modulus, I

y

~

• polar section modulus, • ~

y

y ____

'f

(for round tubes)

'._.-·~

__ x

~~j

",,·V z

Figure P:l Positive directions of axes and an~les on airplane. G(

f3

(alpha) (beta)

S (delta) A €I

(Delta) (theta)

~

(eta)

tI. (lambda) 11' (pi)

I' (gamma) r(mu)

f (rho)

fo 1: (si"ma.)

q, (Phi)

Y(Psi) CAl

(omega)

• angle of attack, de~ees • flight path angle with horizontal, de !n"ge s • nap angle, degrees (elevator, rudder, or aileron): also unit deformation; also deflection. • increment • angle of pitch, degrees, see Figure P:l • propeller efficiency 4/3 1/3 performance parameter • L Lt /L_ 3.1416 s-~ • dihedral an"le, degrees. • Poisson's ratio • mass density of air, slugs/cu. ft., radius of gyration, inches. • f at standard sea level conditions. 0.00236 slugs/ cu. ft. • sum • anl'le of roll, degrees (see figure P:l); also angular deflection • anrle of yaw, degrees (see Figure P:l) • anpular velOCity, radians/second.

• Air Commerce Manual, CAA, U.S. Dept. of Commerce. AN • Army-Navy Standard Specifications. ACIC • Air Corps Information Circular, U.S. Army ALCOA • Aluminum Company of America. ASME • American Society of Mechanical Engineers. CAA • Civil Aeronautics Administration, U. S. Dept. of Commerce. CAR • Civil Air Regulations, CAA, U. S. Dept. of Conmerce. DC • U. S. Dept. of Conmerce. AAFTR • Army Air Force Technical Report JAS • Journal of the Aeronautical Sciences NACA • National Advisory Committee for Aeronautics (U.S.) • Reports and Memoranda, Aeronautical Research Committee (Gr. Prit.). sAE • Society of Automotive Enf,ineers. TA • Technical Aerodynamics (text by same author). Tl[ • Technical Memorandum (N~CA) • Tecp~ical Note (NACA) TN TR • Technical Report (NACA) WR • Wartime Report (NACA) SAllE - Society of Aeronautical Weight Engineers ACI.I

p:s

INTRODUCTION

It Is the purpose of this Introduction to

survey the trends In the recent past In perfor-

mance and use of aircraft as a guide to fore-

casting the types, performance, and uses which

may reasonably be expected in the near future

(1955-1960), to serve the aeronautical design

student and to a lesser extent the industrial

designer, in selecting types to which he may well

devote his efforts.

I:1 Trends In Types of Aircraft and Power

Plants. An aircraft is a vehicle for the trans-

portation of things and people. The principal

types of aircraft in use today (1951*) are air-

planes and helicopters. as shown in Figure 1:1.

Fig. 1:1 Production and licensing data,

U.S. airplanes and helicopters.

Data from CAA Journal, 1953- Auto-

motive Industries November 15,

1953. and Aviation Week March 2,

1953.

A current small (3 seat) production model

helicopter (Hlller) Is shown in Figure 1:2; a

current large helicopter (35 seat Slkorskyj is

shown in Figure 1:3. Note in Figure 1:1 that in

1953 helicopter production (chiefly military)

begaln to approach civil airplane production. The

great boom in helicopter production followed the

discovery in the Korean war of 1951-53, by the

U.S. military services of the enormous military

utility of helicopters. Quoting from an article

in Time magazine on the contributions of Igor

Sikorsky to aeronautical development ("Uncle Igor

and the Chinese Top,"Time November 16, 1953, PP

25-28) "—few machines have so captured the nation-

al imagination. The Marine Corps has long since

Generated on 2013-11-09 15:20 GMT / http://hdl.handle.net/2027/mdp.39015000998867 Public Domain, Google-digitized / http://www.hathitrust.org/access_use#pd-google

adopted the helicopter as Its answer to the atomic

bomb, and proposes to send rotor-topped whirl'y-

blrds hurrying inland from carriers far at sea, to

1:1

establish the beachheads of the future. The Army

has begun supplementing trucks with helicopters,

and in so doing Is regaining a disregard for rough

terrain it has not been able to afford since the

day of the mule. And to-day no naval aviator

leaves a carrier deck without knowing that a hel-

icopter Is hovering nearby, ready to swoop and

pluck him from the sea if he is forced down."

Civilian uses, such as cropdustlng, airmail over

traffic-congested cities, and rescue and construc-

tion work in remote or mountainous country, have

also developed rapidly in the last few years.

Fig. 1:2 Hlller model 12-B 3-seat 200 hp

helicopter in production in 195^.

Courtesy Aviation Week.

Fig. 1:3 Sikorsky model S-56 35 passenger

3800 hp helicopter under test In

1951*. Courtesy Aviation Week.

The helicopter is essentially not a high

speed vehicle (see Chap. 5), and speeds much over

20c mph are not at present in sight except by

c;onvertiplanes (helicopter-airplane combinations,

like Figure 1:4] of dubious economy, but excellent

military potential. Because of the enormous

growth of the helicopter in industry, this 10th

edition of Airplane Design has undergone major re-

vision believed sufficient to Justify change of

title to "Airplane and Helicopter Design."

Fig.1:4 McDonnell XV-1 convertiplane. Time.

INTRODUCTION 1:2 While most military airplanes are turbojet powered in the interest of maximum possible~vel high speed (typical current jet bomber and fighter airplanes are shown in Figures 1:5 and 1:6 respectively) the future of the commercial jet airliner is still (1954) in some doubt partly in view of the unfortunate safety record to date of the British "Comet" jetllner.(Figure 1:10) (Note comparison with other forms of travel in Figure 1:7.)

The piston-engine powered airliner, exemplified by the best-selling DC-7 shown in Figure I:8 appears to be well established for many years to come. 500~-----

I

Fass. Fatalities per

, 200

i r

100 :.:illion pass. :.:iles

~

100t"-.,

scl 20

i L ,

10, i

5~ i

..

~

!

2 •

Fig. I:5

;"utohlo:iles &

~.s . ~'~irlines ~

Boeing B-52 long range jet bomber, key airplane in current plans of the USAF strategic air command. Courtesy Aviation Week.

"

!

.2. I

I

.Il N

0-rl

':'rains-+-

I '-C

N

co N

0

(\J

('1'"\

(""j

C'\O\O\O\

r-!

r-l

Fig. I:7

r-i

r-I

-:J '-.(;

co

0

\

~

N

....::t ...:::; 0-.0-.0\0'\0, r-i r-I rl r-l rl f""""\

0\

('i"\

Safety record of airplane, automobile and train travel. Aviation Week, Feb., 25, 1952, CAB Journal, and National Saf

~~5t-, '-~N ~",":-: .....~

_ ...., ....,-

9tt

taper of 2:1 can be investigated later. For a oonventional tail arrangement, oonsisting of vertioal and horizontal tail, tentatively assume Svert.- 0.075 S and Shoriz. • 0.15 s. For a Vee tail assume a 45 0 dihedral and a total tail area of 0.225 S. This assumes that the Vee tail involves no saving in tail area though there is usually a saving in drag. Locate engine, propeller, pilot, passengers, baggage, and gasoline so that the oenter of gravity (c. g.) will be near the 1/3 ohord point of the wing. Passengers, baggage, and gasoline should be as near to the c.g. as possible to avoid a ohange of c.g. location when these items are absent. Assume a tractor monoplane unless otherwise speCified. The advantage of ohanging to a pusher can be oalculated later. Locate wheels and tail so that the zero lift ohord of the wing may be inclined about 20 0 up for landing and so that the fuselage center line incidence will be about zero at an absolute angle of attack 0( a _ 3 0 to 50. Assume a retractable landing gear; advantage in weif'ht saving may be balanced against drag reduction in a later investigation. (15) Third weight estimate and balance table. Estimate weights of major aiI'plalle oomponents, using charts on pages A5:l to AS:3. Measure location of parts in terms of distance from the nose on the preliminary layout sketch. Prepare table showing o.g. location. (16) Performance Calculation. Revise parasite drag estimate, using 'the data in Appendix A2 and caloulate the specified items of performance. If Appendix A2 is found exoessively condensed, oonsult Technical Aerodynamics, second edition. If the speoified performanoe is approximately fulfilled, the design is ready for an investigation of the effect of changes in design on performance. Suoh an investigaticn involves an evaluation of relative importanoe of decrease in weight (for a given strength) and an inorease in minimum drag. Reasonably accurate evaluations of these faotors is not possible until a speoial study has been made of the relatim between weight, strength, and oost and is therefore deferred to a later chapter. Preliminary comparisons based on average weight and strength of ourrent deSigns are however possible without suoh detailed study and are therefore presented later in this chapter. 1:3 Example of Preliminary Layout. As an example of aPplication of the above method, preliminary layout caloulations will be made for an airplane to meet the speoifications given in article 1:1. These steps are numbered to correspond to the steps outlined in suggested. procedure given above. (1) First estimate of gross weight: For 400 lb. payload, estilliate gross weight W~x 400 • 1600. For small ships like this, the term "payload" may be used to include the pilot (as here), though this usage is not strictly in accord with the NACA definition. On transport ships, the pilot (and erew) are not oonsidered "payload". (2) First estimate of wing area. Sinoe the speoifications Cill for high 11ft fUll span flaps, it will be assumed that max - 3.~.For a stalling speed of 55 mph, the necessary wing area is calculated to be S 1600 O.OO25ElX 3 x (55)2 • 69 sq. ft.

ct

tentatively assume S • 70 sq. ft. as a small change in wing area makes an even smaller ohange in stalling speed. (3) First drat estimate. As suggested in the layout pro-oed.ure, assume thi the arig of tile wing tail, and fuselage are represented ~, a minimum airplane drag ooelfioient of On • 0.012, whioh might be increased to 0.015 on account of the windshield and the fact of the landing wheel is not fully retracted. A 1Il0re detailed drag estimate is to be made later but it may be noted at this time that the minimum drag coefficient of 0.015 here assumed. is le~.

LAYOUT DESIGN OF LI!1HT AIRPLANES

1:5 (b) For an absolute ceiling of H • 10,000 on p. A2f:5, on the next-to-top curve of the absolute ceiling group, read A. 28.

For a wing area S - 55 sq. ft. calculate Om • Sib 55/20.8 • 2.65 ft. and the aspect ratio is ~ • 20.8/2.65 • 7.85. This is a reasonable aspect ratio and will be used in making a layout sketch of the airplane.

The smaller of these two values of A is necessary to fulfill both the ceiling and climb specifications.

(14) Preliminary layout sketch. The wing will be sketched With a 2:1 taper ratio as suggested in the outline of procedure. The root chord will be Or • 2 (2.65) - 3.54 ft.; Ct· 1 (2.65)· 1.77 ft.

(13) calculation of necessary wing span. Using 23 and the cfuii't on page A2f:3 and rhJlm/f _ 54/1 - 54, read LsLt • 86. USing Lt _ 23.5 as previously determined, calculate Ls - 86/23.5 - 3.66 _ w/eb 2 •

YLt •

A •

~

To estimate e note the charts on page A2f:l and observe that ew is usually between 0.8 and 0.9 and that it is somewhat reduced by the effect of fuselage drag variation with angle of attack. Tentatively assume e - 0.8 (to be verified after the aspect ratio has been determined) and calculate

-.J O.B

• 20.8 ft.

1270

!"3'""

These numbers may be rounded off to 42 inches and 21 inches respectively, a5 wing drawing dimensions are usually specified in inches. The span will then be 250 inches. A sketch made in accordance with suggestions in the outline is shOJlll in Figure 1:4. The fuselage width will be made just sufficient to accommodate a single passenger. pafe Al:8 shOirs that, among the airplanes listed, only one airplane has a seat width less than 24 inches. Accordingly the inside fuselage width will be 24 inches at the front seat; the outside will be made 26 inches to allow for fuselage wall thickness. A good fuselage

x 3.66

,,~\-',, I

I

:

I

I

,, I ,I ~-tJ

(;'r - -:. ?::=. - ..... ""

k-------------~--·---ZSO-"---------~

" " ,dJ" ,In

iJ--,,'*'

1\ 1\

\ \

"

II

\I

/ ./

I

\ \

II

,

~

~-

-J0

Fig. 1:4 Preliminary layout sketch for layout example of Art. 1:3. See pp. I:~ and 1:12 tor photogr'aphs of airplanes of th1e type.

01I

-.~

I

LAYOUT DESIGN OF LIGHT AIRPLANES 1:6 shape has its maximum width at the middle of the fuselage: at the middle our fuselage can be about 28 inches; only 21 inches is necessary to submerge the engine completely within the fuselage. The engine Will be located tentatively with ~ts e.g. approximately at the trailing edge of the wing. The passen~er will be located ahead of the main spar which will be assumed at the quarter chord of the wing. A total passenger compartment overall length of 72 inches will be provided on the basis of cabin dimensions on page A8:1. The tail length frem the quarter chord of the wing to the elevator hinge line is made a little over three times the mean chord, in this case 100 inches. The total Vee-tail area (actual not projected) was made about 13 sq. ft. which is 23.7% of the wing area for this sketch. Investigation of stability and control provided by the tail surface may, and usually does, lead to chan~es in the dimensions and/or proportions. A retractable nose wheel is shown to accompany the single main wheel; a retractable tricycle gear should also be considered as an alternate design. (15) Third Weight estimate and balance table. Having the prenminary sketch (Figure 1:4) showinl2' approximate sizes and locations of various parts of the airplane, a more accurate weight estimate and a balance table can be prepared. The form shown in Table 1:1 is suggested. Discussion of the numerical values for the various items follows the table. Table 1:1 Weil2'ht and Balance Table Weight Weight-(empty) Lbs~ 1.

II.

In.

Mcment Lb.-in.

110 16

105 200

11,550 3,200

110 50

100 115

11,000 5,750

------

176 32 5 10

150 150 150 150

26,350 4,800 750 1,500

B. Prope11or-(Complete) C. Lubricating System - - D. Fuel System

41 4 15

135 150 105

5,530 600 1,580

Fixed Equipment, Total-126# Instruments Surface Controls Furnishings & Safety Equip. D. Electrical Equipment - -

10 11 68 37

50 80 70 80

500 880 4,7&J 2,960

70 115

28,000 11,280

Structural, Total-3l6# A. Wing Group B. Tail Group C. Body Group 1. Fuselage - - - - - - 2. Landing Gear Power Plant, Total-283# A. Engine Group 1. Engine-(Dry) 2. Accessories 3. Controls 4. Cowling

III. A. B. C.

Total Weight, Empty Useful Load Pay load and pilot-Fuel and oil - - - Total - - - - - - - Gross Weight - - - -

Arm

725 400 98

""'4%

"Im-

~

e.g. from nose. 123,990 • 101.3

-rm-

Consider the elements listed in Table 1:1 in order. I. Structural elements. A.!Iii First approximation of wing weight can be obtained from Fig. A5:8 on page A5:2. For airplanes of 1270 Ibs. groes weight, read for internally braced wings a unit weight of about 1.75 lbs,lsq. ft.

Check with Fig. A5:7: to determine the design wing load factor, use Fig. A6a:3, calculate W/Bhp • 1270/65 • 19.6 lbs/hp and for W • 1270 lbs. read nIb • 4.4. With the usual factor of safety of 1.5 the design load factor is n' • 1.5 x 4.4 • 6.6 Calculate n'~/1oooCm • (6.6 x 1270) (1000 x 2.65) • 3.16 and read s 1.75 Ibs/sq. ft., which checks rough estimate from Fig. A5:8. page A5:3 can also be used for a good check on wing weight. The total weight of a normal wing (without flaps) would then be 1.75 x 55 • 96 Ibs, but approximately one lb/sq. ft of flap should be added for a double slotted flap of this sort. For 20% chord flaps, the flap area would be 20% of 55 or 11 sq. ft. and the additional weight due to flaps 11 lbs., giving total wing weight of 107 lbs. A few pounds more should be added for retract-able ailerons so the wing weight is estimated at 110 pounds.

ww

Eo Tail Surfaces Fig. 1$: 8 shows that taU surfaces for light airptanes have usually been built for about 1.05 lbs/sq.ft. for airplanes of this size. The Veetail will presumably have to been a little heavier per sq. ft. because of the more severe design loads and a reasonable allowance for the increase in tail surface weight is judged to be 15%, so that the estimated unit tail weight is 1.2 lbs/sq. ft. and the estimated total tail weight is (1.2 Ibs/sJi. ft.) x 13 sq. ft •• 16 lbs. to the nearest whole lb.

C. Body Gro~. In Fig. 1$:9, read fuselage structure weight • 140 Is. For a conventional landing gear of the two main wheels and a tail wheel or skid, read in Fig. A5:9a weight of 75 lbs, this is judged to be reducible to 50 lbs. by using a bicycle landing gear. II. Power Plant Elements. A. Engllle Group. -on-page A4:l, the weight of the Continental A=65 Series 8 engine is given as 176 Ibs, (without hub or starter). Desired accessories are starter and muffler. The smallest starter listed on page A5:27 is given as about 17 lbs. ~1aust manifolding, muffler, and exhaust pipe are not usually listed in weight data, but presumably such equipment could not be much lighter than similar automobile equipment, and would have to weight about 15 lbs. The total weight for engine accessories is thus 17 + 15 • 32 lbs. Engine control units and connections may be estimated at 5 lbs, allowan(".e to provide for' scoops in the side of the fuselage designed for minimum drag and good air flow to the cylinders cannot be carefully estimated at this stage, but a figure of 10 lbs. is believed to be conservative. B. Propeller. The weight of six blade 60" diameter dural prop is estImated from fig. A5: 2 to be about 55 lbs.(extrapolated) and it may be noted that wooden propellers usually weigh 20 or 30 per cent less than dural for the two blade type. Accordingly the propeller weight is here estimated to be .75 x 55 • 41 lbs. C. D. Lubricating and Fuel Systems. Use Figs. A513 and A5:4, nth a 15 gailon gas tank and a one gallon oil tank. For Dural gae tanks and piping read 15 lbs. and for the oil tank read 4 lbs. (both extrapolations slightly off the chart). The total power plant group is the sum of these items or 283 Ibs. III. Fixed Equipment Elements. A. Instruments. pages A5:31,32, and 33 read air speed indicator 0.5 lbs., air speed tubing 0.5 lbs., turn-and-band indicator 1.4 lbs., magnetic compass 1.2 lbs., oil pressure gage 0.3 Ibs., magnetic tachometer 1 lb., oil and air thermometers 1 lb. The dash arrangement might well be copied from recent model automobiles. The total instrument weight is 9.9 lbs. which we will call 10 lbs. in this table, working to the nearest whole pound.

on

LAYOUT DESIGN OF LIGHT AIRPLANES 1:7

B. Surface Controls. SUrface control are usually about 10% of the Wlng we2fht (in which allowance has already been made for weight of flaps and retractable ailerons) so that the weifht of the surface control may be assumed 10% of 110 or 11 lbs. C. Furnishings and SafetyEquipment. For seats refer to page A5:36 and choose the 12ghtest pilot seat listed at about 7 Ibs. plus the li~htest quick detachable parachute at about lB lbs. For flooring, including baggage space floor, allow, 12 sq. ft. of 1/4 inch plywood or equivalent (at o.B Ibs/sq. ft., page A5:8 + 0.2 Ibs/sq. ft. for carpet) to ~et a total of 12 Ibs. for floor material. The only safety equipment item here contemplated would be a 1 quart fire extinguisher listed on page A5:36 at 6 Ibs. The total for furnishings and safety equipment is 68 Ibs.

D. Electrical Equipment. Battery (12 volt, shielded, page $:28) 21 lbs.; two-way radio trarlsmitter-andreceiver, lightest outfit quoted on page A5:35, 16 Ibs. The foregOing figures give a total weight empty of 725 Ibs. Adding the useful load(pilot, passenger, fuel, and oil) of 498 Ibs. as in the previous (record) estimate; the third estimate of "ross weight is 1223 Ibs. Thia new gross weigh estimate is 3 - 1/2% less than the Second weight estimate but this decrease may not actually be realized and accordingly the performance will be calculated on~he basis of the larger figure of 1270 Ibs. It is of course not contemplated that all the specified instruments and equipment should be supplied at the advertised price of the airplane (any more than they would be for an automobile), but the design load should be sufficient to permit maximum eqUipment. Performance should probably be estimated both liFht and heavy. Heavy (maximum permissable load) for the deSigners information. Light for advertiSing, (misleading perhaps, but perhaps ethically justifiable on the grounds that the plane will rarely fly with its maximum load). The balance table and the third weightestimate are combined in Table 1:1. The lever arms of the vdrious prinCipal items were measured on the drawing from the nose of the airplane to the c.g. of the component. The sum of the moments divided by the total weight gives the airplane c.g. distance from the nose. This point was plotted and its distance measured from the leading edge of the mean geometric wing chord and expressed as a fraction of the mean wing chord, because the c.g. location on the mean chord is one of the principal factors determining the longitudinal stability of the airplane. The following rules may help in estimating the c.g. location of component elements: For wings assume e.g. at 40 to 45% of the mean winr chord. For tail surfaces assume c.g. at hinge line. For fuselage structure, assume c.g. at 40 to 45% of the overall length depending on fuselage shape. For engine, the center line of the cylinders is usually close enough. Engine accessories are on the end away from the propeller hub. Engine control system weight is chiefly in the cockpit. Lubricating system is usua'ly as near the engine as convenient oil chanpng will permit. Fuel system weight is chiefly gas tanks which should be located near the 0.3 chord point to avoid shift of c.~. with gas consumption. Instrument weight and surface control weight is chiefly in the cockpit, which moves the c.g. of the whole item towards the rear of the cockpit. Furnishings are in the cockpit; electrical equipment weight is chiefly battery, location of which is optional but puttine the battery far from the en~ine increases the weight of starter cable. Passenger and pilot are in cockpit; caggage should preferably be at c.[:. to permit balance with or nthout baggage, but other locations may be used if baggage is always replaced by ballast, (avoid this if possible). Movable eqUipment, parachutes, fire extinguishers, etc., must be near the cockpit to be of use.

From the Table 1:1 the e.g. location from the nose is calculated from the sum of the moments divided by the gross weight and this point is sho'llI! on the drawing as heavy plus Sign with a circle around it. On the drawing the distance of this point measured from the leading edge of the mean chord is found to be 12 inches and the c.g. location on the mean chord is 38% of om aft of the leading edge of Om. This c.g. location is somewhat farther aft than was intended, but the c.g. location changes as the desirn progresses. The wing may later be given a 11 ttle S1reep forward or sweep back in order to get a better c.g. locatiCll on the mean geometric chord. Sweep has no important aerodynamic effect at Mach numbers less than 0.3, though for high speed airplanes the critical Mach number may be somewhat delayed by a large amount of S1reep back as discussed in Chapter 4. A large S1reepback usually involves a structural penalty, as swept wings must generally weigh more than unswept wings for a given area and span. (16) Performance estimates.

Using the dimensions in

Fig. 114 it is now poss~ble to make a more accurate

estimate of the parasite drag. Such estimates are given in Table 1:2 Arguments ju!!tifying the estillBtes are ~iven after the table. Table 1:2 Wing Profile, 14% thickness (mean) ( 654 Series) 55 sq. ft. Fuselage, Oval Section, max. 8 sq. ft. 0.0600 0.0200 Cockpit canopy, 0.7 sq. ft. 0.0039 Tail surfaces, 13 sq. ft. Propellor support Strut, 3" wide Wheel, 1/3 Exposed, 6.6" wide, 22" dia. Total

CD

0.0050 0.0090 0.0003 0.0009 0.0019 0.0014

~

The above figures lIl!re determined as follows: Wing Root thickness 18%, tip thickness 9%, mean errectiveness about 14%. For cruis~ng at sea level at 165 mph calculate 0.00256 x 165 • 69.7 lbs/sq. ft. and for w/S • 1270 55 • 23.1 Ibs/sq. ft.., the lift coefficient at cruising is ~ • W/Sq • 23.1/69.7 • 0.331. An airfoil should be selected 1Ih1eh will bave a low drllf coefficient at CL • approximately 0.3. The Reynolds number of cruising flight is 9380 x (mpb) x Cm(ft) • 9380 x 165 x 2.65 • 4,100,000. UBe of NACA low drag 6$ or 66 series wing is suggested. A suitable root airfoil would be the NAC! 653-418 for which the ordinates are listed on page A2a:12. Note from the top of page A2al2l that the thicknesB has small effect on the minimum drag and from the top of page A2al22 that a drag coefficient of as 1011' as 0.005 at a Reynolds number of 4.1 million might be expected 11' the leading edge iI! Blllooth.

i.

Fuselage For data on elliptical' fuselages refer to page A2b:2 and note that the incremental drag coefficient A CD for a fuselage added to a high wing _8 0.0130 - 0.0093 • 0.0037 based on a wing area S • 150 Sq. in. and a fueelage frontal area Sw • 9.29 sq. in. The corresponding Talue of£\CDIr • 0.0037 :z: 150/9.29 • 0.060. Referring to page A2bll f values as 1011' as 0.050 are reported for & high wing on & circular fuselage. For the subject airplane the corresponding value of incremental fuselage drag based on wing area is A Co .• 0.060 x 8/55 • 0.0009 and this is the value shown in Table 1:2. The value of 8 sq. ft. frontal area for the subject fuselage was calculated from the width of 2.5 ft., a maximum depth of 4 ft., and an elliptical cross section. The foregoing figures are for a perfectly smooth fuselage and do not include the cockpit canopy. Cockpit Canopy. USe the lowest drag cockpit canopy reported on page A2b:4, which is tbe one desi~ted 1-2, and for which the drag coefficient A Cntr (. .6 CoF) is given OIl page A2b14 all about 0.02 at Jlach number 0.3.

LAYOUT DESIGN OF LIGHT AIRPLANES 1:8 The frontal area of the canopy is here estimated as 0.7 sq. ft. so the incremental drag coefficient based on wing area for this airplane Co - 0.02 x 0.7/55 _ 0.0003. Tail Surfaces. Use the low drag tail surfaces o~ 6500009 section for Which the ordinates are listed on page A2a:12. Re~er to page A2b:9 and note that a value of C])r (for horizontal tail only, NACA combination 312) is reported as low as 0.0039 for an NACA four-digit-section tail surface. For the Vee tail, there will probably be a slight increase over this figure but this is judged to be compensated by the change to low drag airfoil for the tail surface and accordingly this figure will be used for the incremental drag coefficient due to tail surfaces. The corresponding drag increlD3nt based on wing area, for 13 sq. ft. of tail, is, CD - 0.0039 x 13/55 • 0.0009, and this is the value that appears in Table 1:2. Propeller support strut. It is considered that the engine Will he completely submerged in the fuselage with a cooling scoop of negligible drag. The propeller drive is considered to be by a shaft through a cantilever streamlined strut of about three inches maximum thickness and three ft. length. The drag per foot of such a strut is listed on page A2b:ll at 0.295 Ibs/ft. at 100 mph and the junction with the fuselage is conceived of as being the equivalent to one fitting for ·which the drag is 1.75 Ibs. at 100 mph. Accordingly the drag of this propeller support strut and shaft housing is estimated to be 3 x 0.295 + 1.75 - 2.65 Ibs. at 100 mph. The equivalent parasite flap plate area f is 2.65/25.6 _ 0.105 ft. and the corresponding airplane incremental drag coefficient is f/S • 0.105/55 • 0.0019.

en -

Wheel drag. For the main landing wheel, select tentatiw1y a GoOdrich low pressure landing "heel tire of nominal size 6.50-10 and static load rating of 1300 Ibs., for which tIle outside diameter (as listed on page A5: 22) is 21.8 inches and for which the maximum width is 6.6 inches. The frontal area of such a tire is represented approximately by a rectangle 6.6 inches by 22 inches and the exposed area (one third exposed) is 0.33 sq. ft. On page A2b:ll read drag per sq. ft. at 100 mph of 6.0 lb•• so that for this wheel the drag at 100 mph is 2.0 Ibs. The eqUivalent flat plate area is 2.0/25.6 • 0.08 sq. ft. and the incremental drag coefficient is 0.08/55. 0.0014. Total minimum dra,'

Summarizing the drag estimates the

minimUm drag coer icient for the airplane is found to be

0.0185 instead o~ the 0.0150 hoped for in the preliminary design. This drag coefficient is lower than that o~ any military airplane pictured on page A2b:6 (for which data are given aJ page A2b:5) except airplanes 7 (wind tunnel mockup) and 11 (Bell Airacobra) and accordingly the drag estimate may be considered reasonable. The value of f • 0.0185 x 55 • 1.02 sq. ft. may be used to calculate the performance parameters for the charts on pages A2f:2, A2f:4, and A2~:5 as followel Calculate

Lp

• 1270/l.02 _ 1250.

Usin~ the charts

on page A2r:l, For aspect ratio approximately 8 read

Sw • 0.86 and

A

(l/e )~ • 0.78.

Calculate

/J.

(1/e)f.

'"'"!rls 0.78 x 8/55 • 0.111. With l/ew • 1.16 calculate lie • 1.16 + 0.11 • 1.27, or e • 0.79. Calculate Ls. 1270 • 3.69 and "'0"';'.7';-9'-X--:"(2""0-•. .,.8""3...) 2"-Calculats Lt _ 1270

•

0.83 x 65

Lp/Lt.

Calculate

___ . LT.

Calculate -A..

===~

53.0

A2t:4 tor

I r'tf

Calculate LaLt • 86.7

430 tt.!mi ••

= 10.loo

a:a4 oalaulate

-A..

23 read an absolute ceiling at

The specified performance seems to be "ell exceeded except in the matter of climb, where the fulfillment of the specification is marginal since the span was designed for this climb. An increase of span from 20.8 to 25 ft might well be given serious consideration for better climb from hi~h altitude airports. 1:4 Study of Possible Revisions of Layout. The layout (Fig. 1:4) made according to the rUles given in Article 1: 2 must be considered tentative. Before proceeding with stren~th estimates, it is usually desirable to investigate advantages of possible changes. Changes to be considered for this airplane are, (a) change of wing shape to rectangular wing in the interest of simplicity of construction, (b) elimination of flaps in the interest of possible reduction in cost, (c) retraction o~ partly exposed landing wheel as well as nose wheel, (d) possible change to jet power plant. For each ot these changes there are gains and compensating losses to be estimated. The designers problem is quantatively the net gain or lOBS in these specifications. O~

the changes suggested above only the last will

be studied in detail here because this is the only

one that involves a major change in performance and utility. The foregoing perfo~~ce calculation method is not convenient for jet propelled airplanes because such power plants are rated by thrust rather than bhp. The method recommended involves the use of the Schairer-Boeing Chart (P. A4:9) discussed later in this chapter and described fully in Technical AerodynamiCS, Second Edition. The level high speed can however be determined quickly as followe: A turbojet that delivers 1 pound of thrust can be considered to deliver one thrust horsepower at 375 mph because 375 mile pounds per hour (550 ~t. Ibs. per second) _ 1 hp. The level high speed or the layout in Art. 1:3 driven by engine and propeller turned out to be 185 mph which is almost exactly half of the 375 and this level high speed was obtained with 54 thrust hp. The necessary thrust to produce the same level high speed 54 x 375/185 • 109 Ibs. of thrust at sea level. lIhile no turbojet developing this small amount o~ thrust has been developed, two small turbojets are referred to in this text, namely, the Boeing experimental turbojet (ror which the rotor is shown in Fig. 1:7 and which is there reported to deliver 150 Ibs. of thrust for a power plant weit;ht or 85 Ibs.) and the Westinghouse 9-1/2 inch "baby" turbojet delivering 275 Ibs. of thrust for a power plant weight of 140 Ibs. Since the airplane drag varies very nearly as the square or the speed in the high speed range where the lift coefficient is small and the drag coefficient nearly constant, the estimated high speed of the airplane of Fig. 1:4 converted to m e t propulsion is for the Boeing turbojet 185 x 150. 218 mph at aea level and ~ for the We stinghouse "bahT' jet 185 x _[275 • 295 mph at sea level. YW lIhen allowance is made for the drag reduction due to elimination of the propeller supporting strut, retraction of the main landing wheel, and possible substitution of a plexit;las nose for the cockpit canopy, the sea level high speed of the Westinghouse powered light airplane here considered would be over 325 mph, which would be a phenomenal per~ormance compared with any light airplane available today, though the jet nOise might be highly objectionable. At altitude, unlike the efl~ne-propeller powered airplane, the level hi~ speed improves same-

,3rT7T: V "'P' "'Ii. 86 • 7/ 3.77 - 23.0

A· 2a rea' LtO:!..

=10.100/234 :

1270 _ 23.4

"""54

On page A2f:5 for 12,000 feet.

O:!..

what.

LAYOUT DESIGN OF LIGHT AIRPLANES

1:9 For example, at 13,000 ft. where the density ratio 6 • 2/3 (l/lf. 1.5) the thrust developed at constant air speed is 74% of the sea level thrust according to the general thrust variation on page A4:9. For constant angle of attack, the drag is independent of altitude and the speed of flight must increase in proportion to the square root of the density ratio. Accordingly i f the thrust could be kept ccnstant the level high speed at 13,00 ft. would be 325 x -fW • 398 mph, but because of the reduction in thrUllt to 74% of the sea level ;lue, the actual level high speed at 13,000 ft. is 398 x .74. 342 mph at 13,000 ft. While 13,000 ft is probably as high as private pilot would care to go without Q&bin supercharging. An additional gain of 17 mph to about 360 mph level high speed is obtainable at 25,000 ft. Flight at this altitude would have conSiderable advantage in bumpy weather if a simple cabin supercharging device could be built. It is here estimated that sealing the cabin for supercharging providing an air bleed from the turbojet compressor, and providing a cabin pressure regulator could be accomplished in a plane of this size for approximately 25 lbs. which is of the same order of magnitude as the saving in power plant weight. For flight at 25,000 ft. the wing design cannot be considered anywhere near the optimum since the area was determined largely from landing considerations and span from climb considerations using an engine and propeller. At 25,000 ft. the lift coefficient at level high speed is approximately 0.2 whereas the lift coefficient for maximum L/D is about 0.6 determined as follows:

The drag coefficient for this airplane was estimated to be given by the equation CD • .0185 + .0513 CL2 for the configuration shown in Fig. 1:4, or

Co •

0.0150

CL2 for an improved configuration shown on Fig.

0.0.536 1:5.

+

For the minim1lJl drag condition, deSignated by the Subscript ( h ~ Cnl ~ 2 x • 015 • 0.030 and Cu • :ggi~ 0.54. To

1

•

have a lift coefficient of 0.54 at the high speed condition would require multiplying the wing loading by 3 (di tiding the Wing area by 3) and the Wing span would have to be maintained approximately constant in order to maintain the sa.JIl(value of (L/D)I!IIOX • 0.54/0.030 • 18. The corresponding aspect ratio woUld be 3 x 7.85 • 23.5 and the stalling speed would have to increase by the factor to 96 mph. Thill stalling speed and aspect ratio are judged to be impractical but a compromise high wing loading obtained by dividing the wing area by 2, keeping the span constant, and giving an aspect ratio of 15.7 may be considered as a possible reasonable modification of the design of the Fig. 1:5. With this considerably smaller and therefore thinner wing, the wing weight would be multiplied by apprOximately 4 and about 300 lbs. would be added to the gross weight, giving a gross weight of 1570 lbs. A.dded fuel would also be necessary to maintain reasonable range. If the fuel weight 1~ approximately tripled, giving 260 lbs. instead of 90 lbs. for fuel, & re&soo&ble new gross weight is 1740 lbs.

-rr

I I

I

I

I

i'~:

, L

~:.':

:: =:':,\, '-'

I

-+-":'--t--~i-'--'l-~_

_ "'_-i_o----

I

~--------------------~---250'-'--------~

I'" I j;,/ ($f

..... --

-'~

I

I

\1

11

---\

, ' \

U.

",

"

rI

---_----1.......- - - - - - - - - .... "'"-_ _ _ __

Fif' 1:5. Revision of Fig. 1:4 for propulsion by Westin~hou8e 9.5" turbojet. High speed increases from 185 mph to 325 mph at sea level, to 340 mph at 13,000 ft., and to 350 mph at 25,000 ft. With reduced wing area, the hiph speed can be increased to 390 mph at 1),000 ft.

saKvidHiv ihoti jo «oisaa laoin

■•TPt *ed s)ueo

Z ♦ qv»»/£jt deejidn «q

Jo 00021 uox)onpoad jaeX/oooX no ptnq 'x*X)XUI *»

)soo (i)

•V>TT 009 P"T» ou 'a8u«H *j

*W OOO'St 8ufTfao -aqy «e

(SI * 9) 'UT/'1J 09C(*oado WO) q«Xxo jo e)u -ran -p

qda Sxx o) oiT psede onmprm "o

qdH 001 "dqa %$L 'peads Supsptuo *q

qda Jtj paads 2uptX*)S *a

0JJ84 (Z)

•is 5 ft. but should be checked when the hull has been laid out and curves of displacement vs. draft plotted. The beam of the hull was determined from considerations of take off. The hull beam (b) must be such as to give a beam loading within ~he limits of P A2a.22. This requires that C... Do/wbi:::=0.5 tp get a reasonable value of 4/R at the "hump" (R/~ is the effective coefficient of friction on the \>ater; the "hump" is the peak of the graph of R vs. V during ~he take-of£'). Since lift on the wing~ is 0.32 to 0.4 times the weight, assume .6 ~ (1 - 0.3 ) W =:0.9 w 2oqooo Ibs. in this case. Using Do =200 000 and w 64 for sea water, solve for

=

necessary .!!ing~. I'Iitf 64 read p. A2ta4for - = 27 LsLt = 105 and solve for L = 105/24.5 = 4.3. To solve for b from L8 W/eb~ (for a monoplane. first estimate e: using; p.A2taI·est1mate e, • 0.9 for suitably tapered wing; usini P. A2f'.2est1mate ff/S 60/7100 = 0.008 and for unusually well rounded hullsassume curves nearly as good as for rectangular fuselages. (see TA p195) and read e2 = 0.85. so that e 0.9 X 0.85 ~ 0.76 (this checks The beam has been made 20 ft. in the layout to permit wind tunnel tests for ships of this type). Solvhaving the passengers slsep in beds running athwart ing for b. the ship (two 7.5 ft. beds plus aisle). Take-off calb 260 ft. culations should ~ made as in TA pp.297-30Sto verify the ability to take off, but with the power loading This span will be satisfactory except in that it on which this design has been based and a suitable may not give the most econQrnical combination wing beam as determined above there is little doubt about and fuel weights for the specified range. llanugetting off the water. facturers comnonly check this point by calculating The tail surfaces have been shown as 1000 sq.f't. the combined wing and fuel weight for a given rang horizontal ( = 0.14 S) and 500 sq.ft. vertical. Three with various Sp!ll"lS (trial solution). Airline oper vertical t!lils have been shown (as in the DC-4, Fig. ators might even more profitablydetennine the s 9117) because 'wind tunnel tests have shown that a of minimum cost per passenger-mile as proposed in single vertical tail on a large airplane must be ve!"! Chapter mof this text; more span will save fuel, large for adequate stability and control. This should but the first cost of the airplane (and possibly be investigated in a wind tunnel i f possible, otherits hangarage or dockage costs) will be higher; th wise as in TA Chap.1I. best span will depend upon the number of airplane bui1t and used and the schedule of operations of The general arrangement of the interior contemthe airline-items difficult to predetennine. Sinc plated is that the pilots and navigator should be up the calculated minimum span of 260 ft. gives a front for good vision, the engineer and radio operator reasonable aspect ratio of 2602/7100 = 9.5, the (2 men) should have an office up under the wing, and ship will be tent.ative~ drawn up with this span, there should be 4 cabins (forward, center lower, center but it is quite possible that a span of 300 ft. upper, and aft), with 2 stewards and 2 stewardesses would give a more economical ship. and this possi- (total crew = 10 including captain). Interior arrangebility should of course be thourghly investigated ments would have to be wo~ed out in some detail to before the ship is built. determine a desirable seating and sleeping capacity to provide. Comparison with railway accommodations sug(14) Preliminary layout~. This is shown in gests that there would be ample room to sleep 70 passFig.S.IO The !ii!!L has been laid out with a engers in canfort. Further working out of this prostraight center section of 100 ft. span and 35 ft ject is obviously a lengthy task, even if as many men chord (1. = 3500 sq.ft.); the tip sections have are available as can be used. For simplicity in this been made trapezoidal and tapered to 10 ft. at the presentation the third weight estimate will be comtips (S2 = 3600 sq.ft.) to get the specified area bined with the balance table. In practice. the weight of 7100 sq. ft. This procedure is arbitrary but and balance would be checked continuously as the demakes for convenience of engine installation; the sign progressed. assumed taper gives an approximately elliptical wing. lIhich makes for good perfonnance. A root (15) Third weight estimate and balance table. The thickness ratio of 20% (= 7 ft. here) has been weight and balance are tabulated here in Table 1.2 found satisfactory by experience; this can be rebelow in the same manner as in the previous example, duced to 10% or 12% at the tips; best thickness c which should be consulted regarding c.g. lOCAtions of be determined later from economics of design. The the various parts to be scaled from the layout sketch. tail length (e.g. to elev. hinge) has been made Explanations of these weight estimates follow. about 4 times the mean chord (4 x 27.3 OK 109 ft.) to give a long hull with plenty of room. A mean Wing Wj(ght is based on an applied load factor of dihedral of about 40 has been tentatively shown; 2.5 (P. .2) With a factor of safety of 1. 5 the this can be checked later to see whether it corre ponds to the vertical tail size. (13) Calculation

2!

YLt - 1560/24.5 -

=

=

=

=

=

=6.~.3f=

3:6

FLYING BOAT LAYOUT

Grumman Mallard

BRITAIN'S NEW 100-PASSENGER PROPJET BOAT:

Powered by six 5,000-hp. propjet turbines, first of the new Saunders Roe

SR 45 Saro, 120-ton flying boats is on the stocks in Britain. It is expected

to carry 100 passengers and cruise at about 300-mph. fully loaded. Two

decks and sleeping berths in cabins are features of the new type due to

take to the air in 1948. (British Combine photo).

^Courtesy Aviation News, June 23, 1947

Long

Extraordinarily clean lines of the Convair

XP5Y-1 are shown in the head-on view in-

dicating low profile drag. New side view,

taken as the giant flying boat was moved to

the edge of San Diego Bay for launching,

shows details of the remotely-controlled gun

, Glean Lines, and . . .

turrets fore and aft. The unusual hull lines

feature a high length-beam ratio. Bulge «t

the top of the vertical fin houses antenna.

. . . Narrow Beam of the XP5Y-1

while radar equipment is carried in the

plastic nose section. Large size of the Alli-

son T-40 tuiboprops is indicated by the

Courtesy Aviation Week, September 5, 1949 estir

-naximura climb as 3000 rt/min.

empty nacelles on the XP5Y-1. Convair

has installed dummy engines and propellers

to simulate mass and weight of the Allison

5500-hp. turboprop powerplants during

flutter tests. Built for the Navy, the plane is

intended for patrol and anti-submarine duty.

Generated on 2013-11-10 13:05 GMT / http://hdl.handle.net/2027/mdp.39015000998867 Public Domain, Google-digitized / http://www.hathitrust.org/access_use#pd-google

ttes top speed of 390 mph at 25,000 ft 4

3,7

FLYING BOAT LAYOUT

-

-

-

-

_ _ _ _ _.l-~I---

1+-------260 ~ 0 ..~i__---------__401

c::

Fig. 3,10

."

,:~ ';~'-~ ,

Preliminary layout sketch for flying boat of 220,000 Ibs. gross weight.

design load factor is 1.5 x 2.5 = 3.75 = n'. Calculate n'W/1000 c = 3.75 x 220/27.3 = 30 and read wine weight in Fi~AS'7(extended by Driggs equatic as 5.5 lbs. per sq. ft. or 39000 lbs. Assuming the c.g. to be at 1+5% of the mean chord, scale drawing and read 62 ft. from nose to c.g. Tail weight for a ship of this size may well be based on the wing data in Fig .A518; the tail span is 60 ft. and airplanes with 60 ft. span usually weigh 6000 to 10000 Ibs.; read weight about 2 lbs./sq. ft. for horizontal tail; estimate 1.6 lbs./sq.ft. for the three vertical tails. Hence tail weight = 2 x 1000 -t- 1.6 x 500 = 2800 lbs. This is probably conservative by a few hun:ired Ibs. • Hull weight, including floats is based on Fig. AS.9 extended, which is simply about 14% of the gross weight; with larger hulls the structure may be more efficient, but for high speed landings the bottom must be heavy. Hence 220000 x 0.14 = 31000 lbs., which is located at about 40% of the overall length from t.he nose, say 70 ft. ~ weight will be about 1.2 Ibs. per takeoff horsepower (FigA5,l.) (see also Aviation,July 1937, p. 64, an article by A.E. Lombard on "How Many lrigines?") for the twin-row radials in the leading edge or 1.2 x 8000 = 9600 Ibs. Estimate from FigA5.1for the aft engines with radiators and coolant 1.6 lbs./hp = 2400 Ibs. each or·9600 lbs. En~ne aecessories are usually figured at about ~ of the b~re engine weight, here about 700 Ibs. Engine controls for complicated install-

ations of this sort are likely to run as high as 25 Ibs. per 1000 hp en~e, here 200 Ibs. Table 1.4. • an~BManc] illtimate for Flying Boat Group Part shown e t'gh : s. Lever arm Mom. , from nose 1000 ft. Ibs.ft. I. A. Wing structure 60 39,000 2340 B. Tail group 2,800 160 450 C. Hull Structures and i'loats 31,000 2170 70

l

II. A.1. Forward engines, dry

2. Accessories 3. Aft. Engines, radiators, coolant 4. Accessories 5. Engine controls B. Propellers, forward Propellers, aft C. Starting system D. Oil System E. Fuel System III.A. Instruments B. Surface controls C. Furnishings D. Electrical Equip. E. Anchor and hoist

9,600 700

40 45

385 32

9,600 700 200 2,500 2,500 400 900 5,600

70 65 20 88 65 55 60

38

670 45 4 95 220 26 50 340

500 1,000 4,000 3,000 500

30 20 60 60 10

15 20 240 180 5

FLYING BOAT LAYour

3:8 Useful Load. Fuel, J..hooo gallons Oil, 700 gallons Crew, 10 at 170 Ibs. passengers, 50 at 170 Baggage, 50 at 50 Ibs. llovable equipment Yail and express, by difference

78,000 5,300 1,700 13,000 4,000 2,000

60 50 50 60 40

900

80

~

60

5040 265 85 816 160 120 72

~

e.g. is at 12825,000 • 58.3 ft. from nose. This is at 220,000 29% of the mean chord and should be satisfactory. Pas sent;er and baggage requirements ere reduced from prelilllinary estimate when it became evident that there would be no room left in the weight estimate for mail and express. Losr cruiSing speed (say 160 mph) and built-in fuel and oil tanks, would leave room for several thousand Ibs. more useful load.

Propeller weight. For dural propellers with hubs, 17 ft. in diameter, read in Fip. A5:2 a weight of 400 Ibs. for 2 blades; estimate 650 Ibs. each for 3 blades, including constant speed control or 2500 Ibs. for each group of 4 propellers. starting system weight.

As!lUlllS 8 starters (ref. p.

!Sr27 , Item 68 but larger) at 25 Ibs. each; four batteries (. 100 lbs.) and allow 100 lbe. for wiring and conduit.

Oil and fuel system weight estimates require first est~mate of the amount of o~l and fuel needed. The hi~h speed of the ship may be estimated from TOpm/f •

an

8960/141 • 64 and in Fig. A2f:3 read about 200 mph top speed at sea level or 200 (1.08) • 216 mph at 8000 ft. Altitude. Cruising at 75% of Bhpm will give a cruising speed of 216 x 3 0.75 • 194 mph. For a fuel consumption of 0.45 Ibs./bp hr (which is conservative). The fuel consumption per hour is 11,200 x 0.45 • 5050 lbs. and a 3000 mile trip would require 3000/194 • 15.5 hrs. Hence the necessary fuel is 5050 x 15.5 - 78,000 Ibs. (-14,000 gallons). Assume 700 gallons of oil (Fig. A2f:4) at 7.5 lbs./gallon (- 5300 lbs.) and a lubricating system weight of 900 Ibs. per 1000 gallons, or 5600 Ibs. Fixed eqUipment. Instruments are about 3% of the engine weight dry, Surface control allowance is liberal, as serve-controls !light be necessary. Furnishings .at about 60 lbs./passenger will be comfortable if not luxurious. Electrical eqUipment includes an auxiliary power plant, heating, ventilating, lighting, and cooking equipment, as ...11 as ample radio. Movable equiplllSnt as on p. is usually required. Passengers, baggage, mail and express make up the balance. Revision of the weight estimate to carry more pay load and less equipment would be the next step. The performance should then be checked as in the previous layout example. There is, of course, still the question as to how the various parts can be built strong enough without exceeding the estimated night. The development of methods of doing this is the principal subject of the remainder of this text. 3:4 ProblSlll. 1. Dellign a transoceanic flying boat to be powered by four 10,000 hp turboprops and to be assisted in takeoff by two 10,000 lb. thrust Jato (jet assisted take-off) Estimate the lrIIight, balance, pertoI'llance, and payload as in Article 3:3. uni ts of 2 minutes duration.

LAYOUT DESIGN OF HIGH SPEED AIRPLANES AND MISSILES

4il

4>1 - Scope of Airplane and Mlaalle Work. In the last

few years (1949-1954) work on the development of high speed

airplanes and missiles has proceeded at an accelerated and

tremendous pace) many millions of man-hours have been de-

voted to research and development work in thla field. Some

of the results for high speed airplanes are shown in Fig.

4>li for which technical data aro given (in-so-far as secu-

rity regulations permit) on p.A2ft8. Hundreds of supersonic

guided missiles have also been designed, developed, and

tested and a sufficient number have been rejected (or put

into military service, like the Nike) to provide the tabular

information shown on p.A2fi8a. The field of a supersonic

missile or airplane design is not suitable for undergraduate

design projects, as even an abbreviated review of the aero-

dynamic literature involves about 100 pages.*

Note in the tabular data on p.A2fi8 that the research

aircraft shown in Pig. 4:1 are nearly all designed to

operate supersonic as well as subsonic and thet one flight

has been made in an occupied vehicle at Mach 2.5. The

missiles listed on p.A2f16a are mostly designed to operate

in the range from Mach 1.5 to Mach 3.5* At Mach 3-5 missi-

les are well into the *Hest Barrier* which promises to be

more formidable than the 'Sonic Barrier" (at Mach 1) since

thermal stresses as well as cooling problems become involv-

ed in thla region. Only short flights in excess of Mach 4

now appear possible in the troposphere or stratosphere with

equipment on hand or at present conceived; the tabular data

for supersonic flow (ppAli7 to AI1I3) are terminated at

Mach 4* Speeds beyond Mach 4 are now usually considered

'hypersonic*. A substantial amount of research work is

being done in this field.

Mach numbers in the region of 20 to 90 are a matter of

common observation (meteors) usually associated with thermal

dlsentegration of the object (great balls of fire). At Mach

Generated on 2013-11-10 13:05 GMT / http://hdl.handle.net/2027/mdp.39015000998867 Public Domain, Google-digitized / http://www.hathitrust.org/access_use#pd-google

numbera in excess of 35 missiles directed away from the

earth escape the earth's gravitational pull end do not re-

turn to It unleaa propelled. A Mach number of about 25 is

necessary for a missile to become an 'asteroid* revolving

about the earth like the moon in an orbit of its own. The

design of such 'space ships" is beyond the scope of this

text end the following treatment of a simple sample super-

sonic missile is presented only to give an idea of the

scope of the problems involved so that the student who

undertakes a design problem in this field may do so advis-

edly.

4i2. Best L/D at Various Mach Numbers. The maximum

subsonic L/D for a well atreamlined airplane of aspect

ratio 6 or better may be in excess of 15 or 20 but this max-

imum L/D occurs only at reletively high lift coefficient and

correarondingly low level-flight speed for the usual range

of wing loadings. Various studies of wings, and wing fuse-

lage combinations in the supersonic region are represented

in Fig. I|i4 and it la apparent that values of (L/D)max of

better than J are obtainable for complete supersonic air-

plenes up to Mach numbers of 5 or more. The usual turbo-

jet exhaust velocity is of the order of 3,000 to 4,000

ft/sec (M s 3 or 4). Since Jet propulsion is possible only

if the velocity of the raaas of air flowing through the unit

is increased, it is apparent that a turbojet can continue

to provide propulsion only in the low supersonic region.

Afterburners ere usually necessary to break through the

•sonic barrier" between M « 0.9 and M ■ 1.1.

It should be understood that Fig. 4'4 does not repre-

sent the variation of drag of any particular airplane with

Mach number, which is considerably steeper then any of the

lines shown in the auperaonic region. A supersonic air-

plane, like a supersonic sind tunnel throat, must be de-

signed for a particular Mach number and operates very un-

LAYOUT DESIGN O? 'HGH SPEED AIRPLANES

4:2 ~------------------------------------------n.25

.8

.,1

,HI

.7

f"

.6

::;1, ~I

•5

",,'

.t,/ lit>

~

.4

I-

- .5

I

I I I \

While the turbojet with shock-ram intake appears to be the most promising power plant for conmercial supersonic flight, such flight is more certain by rocket power

\ \

~I

"iJ !

in view of the present sta~e of supersonic turbojet development, but the rocket has inherently poor efficiency compared with the turbojet because i t must carry a great weight of oxidizer for the carbQn or hydrogen in the fuel, and the range is thereby divided by a factor of 10 to 20 compared with the turbOjet. Attempts are being made to develop Nuclear Energy Power plants for Aircraft (NEPA division, Fairchild Airplane and Engine Co., pI1.me contractor); if these efforts are successful, the status of rocket propulsion for high speed aircraft may be radically altered.

3

~I (

.3

.2

~i

2000

MPH in stratoSPher

hob 506 600 Fig.

!

The Viles supersonic aircraft shown in Fig. 4:3 apparently contemplates an annular power plant duct entrance around the pilot·s compartment, with a shock wave diffuser being used to advantage to build up pressure in the turbojet inlet, and with an afterburner in the turbojet exhaust (deSCribed in the figure title as an lIathodyd ll ) to supply (at poor efficiency) the necessary thrust to break through the sonic barrier •

edu

!

6ou

30

40

~esSive Heating

1

4:4 Best LID obtainable at various

!

3doo

I

Jlach numbers.

Figure 4:4 represents a plot of the best LID obtainable at each Mach number for the best airplane designed for that lfach number. For Kach numbers between 1.2 and 2.0 the best wing design has a different amount of S1Ieepback for each Mach number as shown by the studies reported on page A2a:34. and designs with no sweepback are quite feasible. Present studies indicate that the advantage of Wing S1Ieepback becomes less as the Jlach number increases. Taking into account the structural handieaps of the S1Ieptback wing, the advantage of neep may disappear entirely at lfach numbers in excess of 3.0. Note in Fig. 4:4 that excessive heating may exist at Kach numbers greater than 2 or 2.5, but this should not discourage the development of such aircraft, as it is quite possible that the heating problem can be advantageously solved, possibly by a friction-heat-regain system inVOlving evaporation and rocket discharge of the coolinp, medi1llll.

The present evidence is to the effect that (L/D) becomes progressively poorer as the Yach number increases beyond about 2.0, though reasonably efficient flight in the region of Kach number 2.0 appears to be quite practical and may have camnercial as well as military applieatiell. An airplane of the size of the p-80R. which can carry sufficient fuel for 2 hours flight at SOC mph, when suitably redeSigned for flight at Jrach lllaber 2, (about 1400 mph, depending on ambient air temperature) could carry 1,000 lbs. of pay load across the Atlantic Ocean (say from Nell' York to London) in a,proximately 2 hours. The camnercial mar1cet for supersonic transatlantic "airmail at the price that 1I'Ould have to be charged to make a profit on the flights remains a matter of conjecture but it is reasonable to stIppose that a market 1I'Ould develop i f the serYice were offered.

4:3 Power Plants for High ape.d AirplaJIes. None of the aerodyilaiilc Studies reported In Fig. 4:li take into account power plant drag or possible regain of drag by use of friction heat.

4:4 Example Layo.ut for Supersonic Airplane. The following preliminary design study of a supersonic sirplane was prepared in the sUIIIIler of 1948 as part of a graduate course in airplane design directed by the author at the Univer si ty of Colorado and submitted as a report by F. J. Kroll, R. C. Maydew, R. J. Naegele, C. v. Osborne, and C. E. Waddell and represented about 400 man hours of study, calculation, and writing. The deeign study was considerably handicapped by the neceSsity of ueing only published data not claSSified as to military secrecy. Fell' it anI dssign stunies of tbis sort bave teen publishen tor military sSClurity reasons, and this study would not be publ1shed i f it were judged to be a design worthy of actual construction. SpeCifications. Design of airplane of 20,000 pounds gross weight, 1,000 pounds payload, and 7,500 pounds ruel load far cruising at Mach number 3 at 60,000 feet altitude. The milit!lI"y and pose1bly commercial si gn1ficanc~ of supersonic aircraft has been effectively pointed out Reference 1, which inclUdes photographs of an experimental rBllljet-pOll' ered air plane wi th jets loea ted on the wing ti ps. Reference 1* pOints out thst Rto achieve an airplane that will have range and carrying ability above Mach 1 is an extremely difficult problem. It must take off and land at a practical speed and ny at first below Mach 1. It IllUst pass through the dangerous transonic band wi thout being thrown out of control or damaged by buffeting. Then it must deal wi th 'he nell' air behavior and enormous drag encountered above yach 1. There is no known design that will do all these things and still be a useful airplane R• Tha size of airplane chosen was arbitrarily cOOsen as approximately the l!I8lI1e groes weight as a DC-3 airplane, mmely 00,000 pounds. SiDee it was realized tlBt the pay_ load would orobably have to be considerably lesll than the DC-3 and yet shruld be a reasonable value, an arbitrary payload at 1,000 pounds was assumed. Sillee it was desired to hava as muoh range as possible an arbitrary fuel weight of 7,500 pounds (2rt'.5% of gross weight) was assumed all reasonable and possible. For military purpoeS8, the ruel weight and gross weight might both be substantially increased in the same layout. All of the information used in making thIs study as presmted here is of nonclassitied 1II111tary status. Lalout Procedure The layout procedure 1I8.S in general that described in Chapts- 1 exoept In so fer as the procedure was necessarily modified by the statement of object and specifioations. A thin sups-sonic airfoil was assumed and a wlng area detet"mined to provide a landing speed of approximatel¥ iOO mph. Tbe wing area, as determIned by a

* Science

Editor; Time YB8Bzlne, Aug. 9, 11148, pp 54-60.

4:3

LAYOUT Dl!SIGN OF HI (E SPEIID AIRPLANES

....

-------, ~

-I

f

,.

i

~

I

I~

~

is

N

I

!!i'

I

~

0 \J\

@

--~

i j Fig. 4: 5

Three-view Drawing Showing Overall DimEil ai ona

'---_~

_________________.1

~

-~---

Fig. 4:&

Drag coetticient (Based on Wing "-rea) VB )lach No.

4:4

LAYOUT DESIGN OF HIGH SPEED AIRPLANES -'-

!

-- -

..

i

~

,

........ ~c-i-:.-· :.(

-~ -''- 1-'-'-

~.. ,\\. "._. .

- -

-

-

.

-.-- .. ~ '

I",

- _.

~i

_.

-_.-

---_('"'l'a

n)

~ "

and the

3'" by a good margin, choose = 0.025, and on the line estimated at D • 36, read w/ • 38 lbs./sq. ft. blade loading, ( R)opt = 360 ft./sec., and Bhp/(W/1000) is equal to 30. Tor alternate solution, follow example at bottom of p.A2e,14 and get (llR) • 330 ft./sec., Bhp • 35. (b) The sea level engine power required, from Eq. (40) is reated to the altitude power, at constant r.p.m" by

(c) To check high speed requirement, read first the sea level hovering power from Fig. 5:8. For D a 40 and ( j . 0.025, read w/ [J • 32, (1l.R)o t .. 330, Ph .. 28.5 h.p. per 1000 poundS~ with a stall limit speed mphLS • 50 (This s~ly says blades will not stall at 50 m.p.h.). To check level high speed attainable with the available S. L. power of 38.2 h.p., calculate Pa/Ph .. 38.2/28.5 • 1.34 and in ~. 5i14, for a medium body drag coefficient f M· • 0.2 (corresponding to f • ~.2 sq. ft. of flat plate for A:: • (fiX 20 ) (0.025) .. 1260 x 0.025 • 31 sq. ft.) read for r • 3 (corresponding to Fig. 5:8), "L· 0.345 • mph L x 1.47, or mph L • (0.345/1.47) x 330 • 1~~ Since the blades stall at 50 m.p.h., the level high speed will be stall limited, not power limited. (d) To check the maximum sea level climb requirement, read in~. 5 :!5, (for P.. /Ph • 1.34, f/Ar[. 0.2, anar· as before) Ch max/33Pb • 0.8, and calculate Ch max • 0.8 x 33 x 28.5 • 730 ft./min. Since only 500 ft./min. was specified, no additional power is needed to satisfy the climb specification.

Items (5), (6), (7), and (8) can proceed on the basis of the decision made in item (4). The feasibility of the foregoing analysis will be contingent chiefly on the rotor weight study involved in the next weight estimate. 1.

5:10. References. Carpenter, Paul J., Effects of Compressibility on the Performance of Two Full-Scale Helicopter Rotors, NACA TN 2277, January 1951.

2.

Glauert, H., Airfoil and Airscrew Theory, Cambridge University Press.

3.

Gessow, Alfred, and Hyers, Garry C., Jr., Aerodynamics of the Helicopter, }~cmillan 1952.

4. Wood, K. D., Technical Aerodynamics, 3rd 3dition. be distributed by University Bookstore, Boulder, Colorado, 1955. 5.

Dommasch, Daniela., Elanents of Propeller and Helicopter Aerodynamics, Pitman 1953.

6.

Slaymaker, S. E., and Lynn, ::tobert R., and Gray, Robin B., Experimental Investigation of transition of a model helicopter rotor from hovering to vertical autorotation, N;'CA TN 2648, Harch 1952.

7.

Slaymaker, S. E., and Gray, Robin B., Power-off flareup tests of a model helicopter rotor in vertical autorotation, NACA TN 2870, January 1953.

8.

Baer, Harry S., Safety shown in autorotative landings, American Aviation, ~ecember 7, 1953.

9.

Sturgis, Raynor F., Factors in Helicopter Economics, Aviation, Nay 1947.

(3. ) Tentativel~ select a P2jer ~8ant of Bhp •

(38.2Io.8~ ~o-l. • to 54 h.p. Consult a list of available engines, such as that on p. A4:1 (or preferably more recent manufacturers data) and note that a 65 h.p. Continental 4 cylinder engine, weighing 170 pounds dry, is the least powerful and lightest airplane engine available. A special two-stroke-cycle engine of 50 to 55 h.p. might perhaps be constructed for the purpose, but tentatively select the Continental A65-8F and see whether the weight requirements can be met.

(4.) Hake second estimate of

~ weight, referring to'fable 5:2 and using average values, note that (power plant weight) • 2.1 x (weight of engine dry) • 2.1 x 170 Ibs. = 355 Ibs. Note also that (weight empty) • 2.5 x (power plant weight) • 2.5 x 355 • 890 Ibs. For 1 hour hovering (the specified endurance is 1 hour; cruising for 1 hour is a less severe requirement as seen in Fig. 5:12), assume fuel consumption of 0.6 Ibs./hp-hr., and Ph • 28.5/0.85 34; fuel needed is 0.6 x 34 • 20 los.; reasonable reserve would dictate a 5 gallon tank with 4 gal. • 25 Ibs. normal take-off fuel weight. Total weight is then \ieight empty· 89.0 Fuel weight • 25 Pilot weight· 200

Gross •

lll,

Ibs.

To

5:11. Notation. Notation used in the foregoing presentation has followed recent NACA practice, with minor exceptions to simplify the writing of the special equations involved in this study. Physical Quantities W R D b

r

p

AO• ,'R2

w.~f2.

A

P

gross weight of helicopter, Ibs. blade radius, feet blade diameter, ft •• 2R number of blades per rotor mass density of air, slugs per cu. ft. sea level air density disc area, sq. ft. corrected disc loading, Ibs. per sq. ft. radius to any point on blade, feet chord at any point on blade, feet mean effective blade chord, ft. ce rotor blade thickness at c e ' ft. rotor solidity ratio, • bcel R rotor moment of inertia, slug-ft. 2 corrected blade loading, Ibs. per sq. ft. approxiw4~e

Sb

a

ce R/2

effective blade area, sq. ft.

LAYOUT DESIGN OF HELICOPTERS

5:16 Physical Quantities (Cont'd.) kr rotor radius of gyration, ft. Velocities

v

~

Aerodynamic Quantities true airspeed of helicopter along flight path, ft./sec. average effective induced velocity behind disc, ft./sec. angular velocity, radians per second tip velocity in hundreds of feet per second forward speed ratio inflow ratio at disc. II· vi/211R free stream l1ach number Reynolds number mean effective blade lift coefficient based on outer half of blade~ For hovering, CLb • 2Lt,/ ~ st(. 75 R)

also such as to tilt the rotor more. Coning and blow-beck of the rotor b18des with forward flight is a stabilizing influence, but most helicopters sre essentially unstable with angle of attack and nothing can be done about it except to provide adequate control. A basic rotor control arrangement is shown in Fig. 5117 involving a 'swash-plate" that ~sn be tilted sbout 2 axes as well as moved up and d01Oll. This srrangement is also appliceble to 3 or more blades. Special devices for control and stabilization of 2 blade rotors ara shown in Figures 5118 and 5.19. A photograph of the hub elements of the Hiller control system is shown in Fig. 5,20. Setisfactory flying qualities of a helicopter require that it have 'apparent" stability both as to stick-position and stick-force; to achieve this it is often necessary to use gyroscopic , hydraulic, spring or electric actuators in the control'system ••

lift on blade, Ibs.

mean effective blade drag coefficient based on outer half of blade ~

Characteristics

T Q

P

Bhp

P-210[/1000

p-* .. r ,.

p/Ph

Phi/Php

f

f* .. 500f

rei'"""

rotor thrust, Ibs. rotor torque, Ibs.-ft. rotor power, ft. Ibs./sec. p/550, horsepower rotor horsepower per 1000 Ibs. gross weight. p .. Pi + Ppb ratio of power in forward flight to power in hovering flight ratio of hovering induced to hovering profile power ratio of induced power in forward flight to induced hovering power ratio of (profile + body) power in forward flight to profile hovering power flat plate area of unity drag coefficient equivalent to the body (or parasite) drag body drag factor rotor thrust C.. T T

coefficien~.

2

(,rrR (12R)2

.--io--..... fAV 2 x 104

rotor torque coefficient. C • Q

Q

e

11

Fig. 5:17BASIC ROTOR DESIGN - CYCLIC CONTROL Courtesy Air Trails" Oct., 195~-~.s-:es

R3 (.il.R)2

P

pAy3xJ.06

MI/#Idd. Jjt.eep

Fig. 5!18

e."" ...

t:tF

YOUNG NFLYBAR" CONTROL

Courtesy Air Trails, Oct. 1953.

Subscripts hovering induced profile body (or parasite) vertical sinldng maximum

empty power plant p~load

engine

Fig.5,19HILLER "ROTOR-MATIC" CVCLlC CONTROL

Courteay Air Trails, Oct., 1953.

5110. Stability and Control Considerations. One of the best published treatments of helicopter stability and control is given in Ref. 5 (p. 5115). Most helicopters are neutrally stable or unstable in both hovering and forward or side~ise flight. but they are controllable and flyable if adequate cycl~ and collective pitch controls are provided. When a hovering rotor diac is tilted there is a sidewise component of thrust or weight which provides e lateral accelerption and soon later a lateral velocity. Even small lateral veloci~ies result in changes in the pitching moment on the rotor tending to tilt the rotor more, and if the body and center of gravity are substantially below the center of the rotor diac (as is common) the drag forces are

Photograph of connection between wobble plate, control rotor~ and main rotor of Hiller 12-B Helicopter. vourtesy Hiller Advertising.

5·17 Table 5.2.

Weight analysis of typice1 helicopters (estimated items designated bye) (chiefly from Ref.9! some items verified by manufacturers; e1so checked against Aviation Week, 25 Feb., 1952).

Manufacturer Model

Sikorsky R6A

Sikorsky S51

Bell

47

Bell

48

Piasecki i'Dl8

Piasecki PD22

Seibel

Kamen

SiiA

K240

~

12-B

~~~~~~~~~~(~S~u~pe~r~s2ed~ed~)~__________________________________________~(A~l~l~e~st~.~)~(Super-

DTII.ENSIONS. :ro\;'ER. PERFOW.ANCE 2 Sests Est. (Pey1oad Pilot )Wp1 400 Reted Ehp, normel 245 Rotor diem.ft ... D 38 3 Blades" b Area, sq. ft. • A 1130 Solidity ratio: " .05e Disc Loading W/A ., wI' /.," 2.3e Blade Loading W/A : w,~ /",,"f 46e Rated 1000 Bhp/W: P 95 Tip speed ,cruisingct1.R)c Level high speed, mph

4 1000 450 49 3 1880 .0465 2.94 63 82 500 103

3 600+ 200 35.2 2 970 .036 2.42 67 85 660 98

8 1600e 600 47.5 2 1770 .040 3.67"91.y 92.y 695 105

2590 574 2016

550 0 1690 3810

2350 850 1500

6500+2500e 4000+

Rotor, total Blades Hub and mechanism

259 159 100

596 376 220

32 9

939

Power plant total Engine, dry Gear transmission Fuel system Starting, cooling, etc.

665 489 176

113 2

559 310 124 28 97

lI'LIGHI'S , lES. Gross weight : W UsefUl loae : Wu Weight empty: We

Tail rotor Body group

~

682

450

~

1D92(

1832

flight controls

WEIGHI' RATIOS For first estimates Gross weight : W Payload ip1 For second estimates Po.er plant total Engine dry weight Weight empty: : ~ Power plant tots1 Wpp

Ws/Bhp Payload/Bhp For later estimates Rotor; Blade wt/A" Hub wt/Bhp See also Fig.A5'5 Transmission wt/Bhp

176 153-

498 4411928 846 397 177 508

6 1200 540 35 3x2 96Cx2 .05e 3.0 60e 94

22 4400 1275 44 31 are. .2 .4 .6 .8 1.0 1.2 1.4 usually given.) See page A7 for current prices. Fig. 61 24 Constituents of nonllal (slowly cooled) carbon steels. (Bullens) Steel is an alloy of iron and carbon and sometimes other elements. Carbon ~ is steel which owes its distinctive properties chiefly to the carbon which it contains. Alloy ~ is steel which TeIIP ., J'o owes its distinctive properties to some element or elements other than carbon. A numerical index sys1700 I--r-,.--+-+--t-+lr..-f tem of specification numbers for steels has been d vised by the Society of Automotive Engineers (SAE). The specification number consists of four of five 1600 t-lIlI:+-+---t-...,..-+-*foo-'" digits (e.g. 4130); the first of the digits indicates the principal alloy element or elements (thus 1500 I---+~+---I--+--.y.-+-oo( "4_ - -" represents a molybdenum steel; index key for other elements given below). The second digit generally indicates the approximate percent of the dominant alloying element, and the last two digits indicate the approximate carbon content in "points" (hundredths of a percent). Thus SAE 4130 steel means molybdenum (4) one per cent (1), 30 points 1300 0 .2 .4 .6 .8 1.0 1.2 1.4 carbon (30) • .§M W 1..2 Element s in Alloy Steel Fig. 6125 Critical temperatures of steels (IronCarbon • • • • • 1 Yolybdenum • • • 4 carbon diagram) (Bullens) Nickel • • • • • 2 Chromium • • •• ; Nickel-chromium 3 Chrome-Vanadium.. 6 Tungsten .7 Silica!.langanese .9

n

6:13 STRUCTmL~

150

100 75

.4 .4

1.2

Fig. 6126 Properties of normal carbon steels (Upton).

1600 Temp. FO

The probable actual yield point anu ultimate strength of air-cooled SAE 4130 steel are seen on p. to be 60 and 95 thousand Ibs. per sq. in.; heat treated as spe cified, the yield and ultimate may be 110 and 130 thousand. Stresses allowed for design ( A6bll} are 50 omd 90 thousand as received from tube mill. for welded SAE 4130 fuselages, not heat treated after welding, Y.P. and Ult. are limited to 45 ~d 80 thousand Ibs. per sq. in. High chromiUlll (llstainless ll ) steels have also been used for airplane construction, the most widely used steel for this purpose being "18-8" (USN Spec. CRS-l). The properties of this steel are given on p.A6bIJ.A high degree of resistance to corrosion can be obtained but at a considerable penalty in structural weight. Some recently constructed stainless steel wings compare very favorably with dural wings in cost and strength-weight ratio. See ANC-5 for a more dtailed specification of corrosion resistant s~eels, particularly CRS-l, ANC-S. See NACA Rept. 615 for a comparison of CRS tubes with X-4130. ~ Properties of Aircraft &YmiMany modern airplanes are made largely of aluminum alloys, chiefly 24/fJr dural. Airplane structures made of dural are lighter than equivalent structures made of SAE 4130 steel, but probably not lighter than structures made of heat treated SAE 6130 steel. The current price of dural is about 3 times that of SAE 4130 steel (See page A712), but still low enough to permit competition of dural with steel in transport airplanes. The price of dural in the U.S.A. appears to be set by the Aluminum Company of America at such a figure as to yield the maximum net return and, unlike the price of steel, may bear no direct relationship to the minim~~ cost of production. ft~ outline of the processes in the manufacture of aluminum and dural follows.

6: 15 Preparation

1400

.m!ll! Alloys.

Av ~TfNtTf

1200 ~ ~

I

1000 f-~ I

j/

~

~;I

~

TFlOUTlr~~ +-I\.... nHI1"£ t--TRoO.T'T£

~

"'.4",rtr",sIT(.

600

CONSIDERA~IONS

In the form Qf steel tubes, this material is sometimes used in making welded fuselages, but the chroTiummolybdenum steel SAE 4130 is more often used for fuselages because higher stresses are allowed and the structure can be made lighter.

1000 Lbe/eq. n.

125

800

DESIGN

' FO/sec .1 I at Cooling rate, 1328 0 F 100 200 300 400

\: ~

~

~l

500

Aluminum is made by passing an electric current through a solution of aluminaa (Al203); Alumina is preFig. 6127 Effect of cooling rate on structure of pared from the mineral Bauxite (approx. 60% AlOH, 25% heat treated steel of 0.45 carbon FeO, and 15% Si0 2 ) by crushing, heating (calcination), (Bullens). grinding, screening, and dissolving in NaOH (Bayer precess), and treating with AlOH to precepitate A1 203 (refThe effect of alloy elements such as nickel is erence S. Mortimer, Aluminum. Isaac Pitm~ Sons, 1919). to slow up the A-M-T-S change, so that an air cool- Aluminum may also he prepared from Cryolite (AlF and NaF) ed alloy steel has much the s~e properties as a but abundant deposits of Cryolite have been found only water quenched carbon steel. Alloy steels are oft- in Greenland and in the Ural (U.S.S.R.). Aluminum is en oil quenched and tempered and the physical pro- found in almost everybody's back yard in the ferm of perties resulting from sucn heat treatment are pre- Kaolin, chief constituent of common clay (A1203; 2Si0 ; 2 dictable with excellent accuracy. 2H20), but the cost of producing aluminum from Kaolin. is at present high. Properties of carbon and alloy steels of nUl!lerous SAE specifications are given in ANC-5, data from Bauxite is abundant in the U.S.A.; deposits are which are summarized on' p.A6bll.Properties and known to be available at the locations shown in Fig. costs of steels and other materials are given on p. 6,28Annual imports of Bauxite to the U.S. amount to aA6b,2.The carbon steel chiefly used in airplane bout 600,000 tons in 1940. Many of the more accessible construction is the mila carbon steel SAE 1025. In Bauxite deposits in the U.S. hav~ been purchased or are sheet form, this steel is made into fittings by cutt controlled by the Aluminum CQ~pany of America, apparently ing with a jig saw or nibbling machine, drillinf, on with a view to discouraging competition, but a careful a drill press, and bending cold over forms. The geological exploration would probably lead to discovery only heat treatment specified in ANC-5 is annealing of additional deposits on abandoned farm land that could (heating above the lower critical temperature and be purchased for $lOO/acre, and Bauxite can be delivered to slowly cooling to remove stresses due to cold working). The yield point and ultimate strength of such steel are given as 25 and 55 thousand Ibs/sq. in. according to ANC-5.

6:14

STRUCTURAL DESIGN CONSIDERATIONS years, the stronger alloy 24-ST has been re~lacing 17ST in airplane construction. The current p~ice of dural sheet is about 50¢/lb. (See p. A71 2) in the thicker sheets. The 24-ST alloy has a Y.P. about 20% higter than l7-ST, but there is little, i f any, gain modulus of elasticity. The price difference is about 10%; an airplane built of 24-ST instead of, 17-ST is likely to be lighter and no more expensive. A pictorial outline of the production of aluminum from advertising of· the Aluminum Co. of America is given in

page 719.

Figure 6128 Bauxite deposits in the U.S. known 0, now being worked X. Locations of abundant cheap electric power shown thus

e.

Numerous other alloys are produced by the Aluminum Company of America for special purposes; the composition and properties of a number of these are listed onA6ba3. The meaning of the symbols designati11.g the alloys is as follows; The wrought alloys are desienated by a number and the letter S; the number denoting tne chemical composition. The most widely used wrought alloys are 3S,. l7S, 24S,24S, and 51S. 3S is used where high strength is relatively unimportant and welding is necessary, as in gas tnnks and chairs. l7S is the alloy usually referred to as dural;

the mill for $30.00 per ton of aluminum bi1lets produced. Operation of the crtlshins machinery, vsts, and furnaces is several times as expensive 0.S iron per ton of product, but need not exceed ~90.0l per ton of aluminum, making ~120.00 per ton (6¢/L. a possible price. The current price of about l7¢ per lb. for No. 1 Virgin 98-99% pure alu~num (See p. A7a2) has made possible a brge 1lJr.ount of expansion and development work in the aluminum industry. The market price of scrap cast alu~nun has l'anged from 9; to 10¢ per lb. for 1941. One of the types of furnace used in the production of alu.~num is shown in Fig. 6129-

24S is an improvement on 17':; '"lith slightly higher strength but lower shock resistance. 25S and 51S are approximately equivalent to 17S in strength when hard, but weaker when soft.

o designates soft or annealed temper after air cooling from about 9600 F. Wdesignates the condition after water quenching from 9600 F. T designates the condition after aging. Time and temperature for aging are sho...n below for several1.lloys.

Aging Temp. Aging Time (ALCCA alloy l7-ST) is made by adding Roan 4 days small quantities of copper (4%) and magnesi~~ (0.5 8-15 hrs. 290oF. t .. the molten aluminu,.... The dura.l is extrude4 from 18 hrs. 3l5 0 F. the furnace in a pla~tic state and must be care! :::ooled to get the strength specification p.A6b13 Thus 17SW becomes l7ST after 4 days at room temperature, Sheet dural ImlSt be heat treated after cold working more than 90% of the chan~e having taken place in the to retain its high strength. \'iith:l.n the last few first day. Dur81 rivets are usually kept in powdered solid CO2 ("dry ice") after quenching until a few minutes before they are used to keep them from hardening. Durn,l

H designates the condition produced by cold working (strain hardening) of the 0 temper material. The skndard tempers 1/4H, 1/2H, and 3/4H provide a gradation of properties. On A6b. 3 the symbol Al has been used to deSignate "Alclad" alloys, which have a thin coating of electrolytically deposited pure aluminum to reduce corrosion. RT designates material that has been rolled after heat treatment to give still higher strength, ductility is of course sacrificed.

Fig. 6129.

Heroult type fUrl)B.ce.

Under the action of salt water (and air with salt spray ) dural corrodes ne~rly as rapidly 8S steel ., though aluminum "rust" is white instead of brown like i ron rust). and the corrosion while under stress is much more rapia than unstressed corrosion. The effect of the stress-corrosion is particularly deteriment:ll to the fatigue strength. A coating of pUT'e aluminum, electrolytically deposlted, 1s sometimes used to reduce corrosion, though thf' "Alclad" dural is little better than painted dural in resistance to stress corrl'lsion.

srRtJCrtRAL DESIGN CONSIDERATIONS .1. Comparison of Materials. A quantitative comparison of materials to determine the most desirable material for any particular airplane part of Bub-assembly is possible only if the airplane specification includes a careful statement of what features of the complete sirplane are considered desirable. An implication of meat specifications is that psyload-miles per dollar is a measure of the desirability of a particular material or type of eonstruction as compared with another material or type of construction for an airplane of a given power plant and external dimensions. Some of the basic data for a comparison of wood, dural, and ~ for use in airplanes are given in Table-b75. Such a comparison necessarily involves numerous arbitrary assumptions, but msy be interesting partly because it shows what assumptions are necessary. The wood chosen for the comparison is spruce of 15% moisture, because thia is one of the best end most comffionly used woods. The dural chosen is the ALCOA alloy 24-sr because it is one of the best of the str aluminum alloys and is superior to the newer and more widely used 75S at temperatures over 3000 F. as shown in Fig. 6:9. Two steels are chosen because both kinds are widely used and a study of either alone would not give a fair comparison: SAE 4130 ('chrome-moly') is the typical steel tube material used in fuselages; heat treated SAE 3435 (Chromenickel) is t)~ical of good British practice in steel construct ion. A structural weight modulus is developed in Table 6:5 by compering stiffness end strength with weight. Tension and bending members in an airplane are designed on the basis of yield point strength, but compression members are designed pertly by stiffness (Euler column formula applicable to long columns involvas only stiffness, not strength).

The weighting of the strength/weight and stiffness/weight ratios is arbitrary but believed to be approximately eorrect. Table 615 shows that a good glass-plastic will probably give the lightest structure, air cooled 4130 steel the heaViest, but wood and dural structures are nearly as light as heat treated steel. A structural cost modulus may be developed by multiplying the structural weight modulus by estimated mat,erial and labor costs per pound. Materiel costs include allowance for typical waste of material. Direct labor costs imply an assumed method of construction. For t he wood structure, plywood covering is essumed to be necessary to get ample stiffness with as clean a design as would be possible with metal structures. For the dural structure, sheet metel torsional bracing is assumed necessary; for the steel structures torque brecing of wings like the Junkers end Lorraine-Hanriot bracing (Figs. 818, ellO) is implied. Note in Teble 615 that with current high labor coste in U.S. factories (average over $2.00/hr. in 1954), material costs becane a relatively minor factor, so that even very expensive materials like titanium can also be considered for commercial airplanes. Titanium structures will often be little heavier then aluminum, and the completed structure msy cost only twice as much. Labor costs are always difficult to estimate and are subject to enormous possible improvement. In effect Table 6'5 expresses the judgement that it is possible to build glass-plastic airplane structures for about half the cost of steel, dural, or wood. The structure cost modulus is, of course, something that varies from day to day end place to place. The desired cost moduli for calculating pe yload-miles per dollar fluctuate eTen more widely. The development of cost moduli for deSign is considered in the next chapter,

TABLE 615· Comparison of Materials For Ai!:plane Use Materiel Specification D = DenSity, #/1n.3 E = Elasticity Mod. 10 6 #/in.2 Y = Y.P., .002 set, 1000 #~in.2 U = Ult. Stress 1000 #/in. Rl • Stiffness/weight (E!D) R2 Strength/weight (Y!D) R • Rl • R2 Structural Wt. Modulus. 1000/R Materiel Cost, 11/53, $/lb. Fig. A717 Min. direct labor cost, 11/53, $/lb.,est. Structural Cost Modulus' $

=

Spruce 15% moisture .0156

1.3 6.2 9.4 83 396 479 2.10 .23 15.00 32

Steel SAE 4130 .284 29.0 60.0 95. 102 211 313 3.2 .20 10.00 33

SAE 3435 130. 155. 457 559 1.8 .40 est. 20.00 37

Dural 24-sr .101 10.5 40.0

62. 104 398 502 2.0 .55 15.00 31

Glass-Plast ic 50% glass by Vol. 143 Fab.-114 Finish .067 5·0 50 est.

n

75 750 825 1.2 .85 10.00 13

6:16.1~. Solution of problems like the following is a prerequisite to becaning a successful manufacturer of airplanes, or even to profitably building an airplane of original design. No single chapter or book cen give the information necessary for en accurate answer, since economic, social, and political forecasting are as necessary as knowledge of manufacturing end management technique. The practical man calls it 'experience', but assistance to an accurate answer can be found in the manner end from the sources shown in the foregoing psges.

1. Estimate the cost of production end delivery date of en airplane similar to Fig. 116 on which it is proposec to start ordering material May 1, 1955. State assumptions necessary to arrive at your estimate.

6:16 STRUCTURAL DESIGN CONSIDERATIONS

6,17 stress Analysis The strength of fNery member and fitting of an ,..irplane intended for commercial license must bt; calculated by the manufacturer with a view to demon st ratin,;' to the CAA that no member 'NHl fail (or suffer permanent deformation large enough to impair the usefulness of ~he airplane) under a~v of the loads vmich are likely to be applied to it. Methods of estimating the loads which may reasonably be expected in servi~e have this ehapter. The been discussed in current chapter deals flith (a) methods of calculating the unit stresses in the members due to the applied loads, and (b) calculation of margin of safety fro:n the applied unit stresses and the known unit strengths of the mcterials. This is the customary procedure even tho the actual failure may be due to instability (as in long columns) rather than to excessive stresses. It of course does not follow that calculations showing a positive margin of safety will assure a satisfactory design. A design may be strong enough and still be unsatisfactory because of lack of rigidity or durability against fatigue failure. In fact, it i~ frequently stated that if fatigue, rigidity, and shock' resistance are used as criteria for design, the structure will nearly always be amply strong, and strength calculations are unnecessary or unimporbnt for design. Strength calculations are required, however, and probably quite rightly, before a license will be granted. Fatigue strength and rigidity are more difficult 'of predetermination, and suitability of a design in these respects is c~1only left for determination by flight tests. The student is assumed to be familiar with suo elementary te.~books on Strength of Materials as those of Prof. A.P. Poorman of Purdue or Prot. J. B. Boyd of Ohio &tate. Some ot the following mater ial constitutes a v~ry oondunsed review of the elementary texts; the rellll1nder is in the nature of an abstraot of speoialized books on airplane stress analysis such· as' "Airplane Structures" by Niles and Newell, "Airplane stress Analysis by Pro!. A. Klemin, or "Analysis and Design of Airplane Structures"by Prof. E. F. Bruhn, but aims to include also sane of the concepts presented in the more recent advanced work ot Prot. S. Timoshenko. Cognizance is also taken of the developments presented in ANC-5 (Gov't. Printing Office, 25¢) which should be studied in conjunction with this Chapter. The student who proposes to speCialize in stress analysis should of course make a canprehensive study ot the original works. The following material is such as is necessary to a designer in making preliminary stress analysiS, and also serves as an introduction to the speCialized work ot the stress analyst. 6&18 ~.2£ Loading ~ Airnlane Members. Structural ;nembers of airplanes are subjected to three prinCipal tJ~es of loadings, tension,comPfession, and bending. For example, in Fig.6,30(a which shows the forces acting on an air,llane in flight, the member BD is in tension, member CD is in compression, and the portion AB ot the wing is in ~ bendinf. The porti',n Be of the wing is su jected to both bending and cqnpression, aa sholltl in Fig.6130(b). llembors or parts of an airplane are also so:netillles subjected to shear, as shawn in Fig. 61310r to torsion (t.wisting)as shown in Fig. 6132 Bending nearly always involves shear forces (8 in Fig.6a30b) as well as tension and compression.

Torsion loads often exist concurrently with tension, com;)ression, or bending loads.

Bending and compression

~" tt;'rlmll "~lIIfft~ Bending

~t

Tens i on (b)

Fig.

D

Compression

rs '0.-_-.

6130 Forces acting on airplane and airplane members.

-t::J 1__ [........1........._ _1-

~ (Shear Fig. 6a31Riveted joint and shear force on rivet.

Fig. 0132 Torsion forces on wing of airplane in dive. Stress analysis computations involve (1) the determination of the kind and amount of load applied on each ;nember for each condition of flight or landing, (2) the calculation of unit stresses resulting from the ap~lied loads, and (3) determination ot margin of safety by comparison of applied stresses with allowable stresses. Stress Analysis calculations involving the combination of bending and/or twisting loads with tension or compression loads are more difficult and uncertain because the nature of the failure with such combinations has not yet been thoroughly investigated, particularly if the sections are thin and the failure partly due to instability. 6s19 Calculation ~ l&W ill StaticallY !lIU.tminate Trusses. If an airplane structure contain. just enough members to maintain its shape (and no extra members) the force acting on each member can be determined from the forces acting on the entire structure by means of the principles ~ statics, and the structure is said to be statically detenningtt. If there are extra ("redundant") members, the dist.ribution of forces d8f)ends on the rigidity or the members; a solution which takes acoount ot rigidity and deformation of members in determining the loads on the members is described as a method or "consistent defomations" j th~ mathematical treatment or problems involving redundant members cOIlIIlonly uses the d~nonstrablaprop~sition that the actual defo~ ations will absorb less energy than any other possible defonnations, and is hence called the "method of least work". Discussion of statical.ly indeterminate

6:17

STRUCTURAL DESIGN CONSIDERATIONS stru~tures must be delayed until stresses and deformations in the various kinds of members have be8'l considered. For statically detenninate structures, the procedure necessary to solve for the forces in each manhar is described in any textbook of engineering mecha,nics. The treatment of Prof. E.H. Wood of Cornell (Textbook of lIechanics) is particularly recomnended, because no other textbook emphasizes so well the importance of a free body sketch.

tan':~t ;)Ie t .-l

tr.n~\~. _.>

e

Fig.6a34. Free Body Sketch of Pin at joint e. !.Fx

=- l:J.ce

~ Fy ...

1

-

5 c,e _

5

lJt. de 25

=0

L de - 1000

=0

25

=

=

Solve for ce 1200# and de -1000#. The loads in the members were arbitrarily assumed to be tension; negative signs indicate compression. Similar calculations at joints a,b,c, and d would give the load in every member of the truss.

R2

Fig 51 33 Free body sketch of simplified fuselage structure. For example, consider the simplified fuselage structure shown in Fig. 61 33. The specified applied loads were presumably determined by distributing the weights of various parts of the plane to the panel points, b,c, and d, by the prinCiple of moments ("Moment of the resultant of a series of forcp. about any points equals the sum of' the moments of the component forces about that point") and multiplyjng them by a limit bnd.lng load factor. The problem is to find th e forces applied to the variws members of the fusel~ge when tpe specified forces are applied to the airplane. The first step in the solution is to find the reactions Rl and R2' The basic equations are the mathematical conditions for equilibrium, the free body being assumed to be in equilibrium under the action of the external forces and the "inertia forces" (reversed resultant forces which are the forces shown on the free body sketch. I f the airplane weighed 1400 Ibs. and the lLnit. load factor were 5, the total l~t load would be 5 x 1400 7000 lbs. and might be distributed as shown. Of this 7,000 Ibs., 1400 Ibs. is external force (weight); the remainder is inertia force. equations are:

A desirable method of cheCking the calculations by the method of joints is to cut the truss by a line at some point where the line will cut 3 members, and use the portion to the right or left of this section as a free body. Thu8 for the truss of Fig.6.33 a line could be drawn thru members bc,bd, and ad, giving the free body sho1'l1 in Fig. 6135. Writing the equilibrium conditions ~

F - O,~F ;r 0, and~ll7, x 1 '

0 (sum of y - components of forces is zero) Z Fx _ 0 (" " x _ If If ,,) ~ llz - 0 (" "moments of forces about any z-axis is zero)

"It

In Fig.6133there are no x-forces with axes as shown; the equations for finding Rl and ~ are therefore ~F - RI + R2 - 2,000 - 4,000 - 1,000 =- 0 1i1zRl. - 8R2 - 4 x 2,000 - 8 x 4,000 - 12 x 1,000 - 0

=

Solve these equations for R2 - 6,500 Ibs, Rl 500 lbs. To find the load carried by each member, consider the pin at each panel point (a,b,c,d,e) as a free body acted on by the load in each of the adjoining members and by whatever external loads are applied to the truss at that point, and write th" equations which ~~ress the condition of equilibrium. Thus the pin at e (Fig6a33) may be represented as a free body as shown in Fig.6s 34 (This method is described by Niles and Newell as the "method of joints"). For equilibrium of joint e

=0

permits solving lor the three unknown forces.

The

unknown forces can also be determined from moment equations 0, ! lizd = equations moments". graphical

about three z-axes {e.g.1M~a - O~Mzb = 0); ccxnplete solution of a truss by such is sometimes described as the "method of See also Chapter 11 for description of method of solution. 4000#

=

Z Fy -

1/

1000if

bd __

r :d Fig.6135

e ~..-

Free body sketch of portion cde of Fig.6.33

6&20 Unit ~ and Unit Strength. Unit stress is load divided by area. There are basically only two distinct kinds of ~tpstress. tension and shear.

Fig.6136 Block subjected to tensile stress in one direction.

Fig.6a37Block subjecced to shear s~ress in One plane.

6:18 STRUC~JRAL

DESIGN CONSIDERATIONS

, --

lnr T

Fig.6,41 Primary tension and shear ~1 '!. 6.

ft,

38 Block subj ected to 3 tensile stresses an

3 shear stresses.

----fa,A

These are shovm in their simplest form in FigsSa 3 and 6 a 37. The unit tensile stress is ft - PiA, and the unit shear stress is fs - s/A; unit stres~es are generally expressed in lbs. per sq.in. in the U.S. and Great Britain. It is sho'l'l1l in elementary texts that: (a) Tension loads produce a primary tensile stress as in Fig. 6 a36. (b) Bending loads produce she~r stress as we as tensile and compresive stress. (c) Compression loads on long columns produce bending ~s well as compressive stresses. (For short columns, compression may be re garded as negative tension) (d) Twisting loads on circular sections produce pure shear stress as in Fig. 6137 • en sections of other shape, twisting loads also produce tensile &~d compressive stresses. (e) Combinations of bending' and twisting may in general produce three tensile stresses at right angles to each other an shearing stresses in three planes at righ angles to each other. This is the most general type of stress and is shown in Fi 6138 Solution of problems involving tensile stresses in 3 directions and shear stresses about 3 axes is usually not considered necessary in airpl~e stress an sis. Even the simplest stresses (Fig.6a36and6a37) are not ~ure. Prim~ry tension results in second~ry shear, as shown in Fig.6a39, ilnd !"lrimHry shea results in secondary tension, as sho'l'll'l in Fig.6a From these free body sketches it can be stown (Ref. Poorman pp. 41-44) that the maximum secondary stresses are Max. f~ Max. ft

= ft/2 il,t = f s at 0

0 - 450 ,;.nd = 45°

--:r ~YVlA' f~A'

Fig. 5139 Free boay :;k.::t.cn .;;howing secondary shea due to tension.

Fig .6a4O

Free body sketch showing secor.dary tension due to shear.

fsx'11~' qA\ Fig.6142 Free body sketch for finding secondary shear and tensile stresses. Parts of airplane structures (e.g. those subjecte, to bending) are commonly fPl\:liected to both primar:, tension and primary shear as shown in Fig. 6141. Using the free body sketch shown in Fig.6a42, it can be bho~n Ref. Poonman, pp. 216-219) that Max. rS (ft/2) + f~) at Q = 1/2 tan-l (ft /2f s ) Max

ft

= ft/2 +

-..j {ft/2)2 ... f~

at 0 = 1/2 tan-l

(-2fs/ft ) Calculation of secondary stresses is important because they frequently determine whether a member will suffer permanent deformation or fail~re. ~~I

~

I

1

16'1

:°1 I~,- ;~:r:El ~ l' I

i

.

""B!

11--

'.,'"

__ t I __,(-1__

Fig.6143 Mechanism of permanent deformcltion (From Upton) • Permanent deformation is always ~ .!u: shear. The mechanism of ,)eI"TInnent deformation is illuBerated by Fig.6a43. 'Maximum unit shearin'" stress, whether primary or second,,!,"!, is the critenon of permanent defoz:mation. Failure may be I:.lUsed by either shear or tension. Since the U.S. Civil Aeronautics~d now specifies that yield point shall be a b~sis for design of airplane members, shear stresses and shear strenb~h are criteria for design. (Buckling is also a criterion for ~1ny airplane parts; desif,Il loads for buckling are determined not by stress but by stiffness). Strength ot air,Jlane materbls is usually determined chiefly from a tension~. The standard size of test piece usually used is shov.n in Fig.6144, and the results of tyoical tension tests of steel (SAE 4130) and dur~ (17-sr) are shown in Fig.oa4S and Fig.O&46 The term "yield point" is used to refer to the first [Joint where the stress-strain graph becomes hori?ontal, as point Y on Fig.oa45. Note that the stress-strain graph for dural never bec~_es horizontal; for dural it is therefore common prllctice to pick out a point "here the permanent defor;n:>tion has reached an arbitrary critical value of 0.002 as a measure of its strength without excessive deformation. The unit tensile stress at this

6:1~

STRUCTURAL DESIGN CONSIDERATIONS point is commonly referred to as the yield stress of the material. From the foregoing argument, note that the yield point strength (Fty ) in a tension test of steel or dural is reached when the unit s;le:J.ring stress (fl s) on a plane at 45 0 to the direction of the load reaches the critical shearing stress Fso - Ft/2 (see pp. P-3 to P-4 for list of structural symbols). This same critical shearing stress is the criterion for permanent deformation of an airplane member subjected to shear or to combim'!d loads in a single plane.

r- ._- - 2 til -- 100. With practical accuracy, it is satisfactory to omit the second 1000. term of equation (6 a16) for tubes of D/t ColuDn formulas are also used to determine the allowable unit com-oressive load in the compression fl~nge of beams (using L as the dist~ce between points of lateral support), and the allowable unit shearing stress in the webs of beams (using L as the distance between flanges).

tal pressure upstream of shock 1Iave static pressure ratio across shcok wave denSity ratio across shock wave temperature ratio across shock wave local speed of sound ratio across shock wave ratio of total head downstream of shock wave to total head upstream ratio of static pressure to total pressure downstream of shock 1Iave ratio of static pressure downstream to total pressure upstream of wave ratio of velocity (corresponding to Ill) to the speed of sound ....here V == a ratio of velocity (corresponding to Ml) tothe speed of sound r.Jere V =0 ratio of veloci tv (corresponding to 1110 to the velocity where p. =T =0

Pl/po

PVPl

f>2/fJ.

TVTJ.

&2/al

pYPo

P2/P3

P2/P o

v,/a*

V,/a o

v./V

1.00 1.10 1.20 1.30 1.40 1.50 1.60 1.70 l.60 1.90

1.0000 .9118 .8422 .7860 .7397 .7011 .6684 .6405 .6165 .5956

.5283 .4684 .4124 .3609 .3142 .2724 .2353 .2026 .1740 .1492

1.000 1.245 1.513 1.805 2.120 2.458 2.820 3.205 3.613 4.045

1.000 1.169 1.342 1.516 1.690 1.862 2.032 2.198 2.359 2.516

1.000 1.065 1.128 1.191 1.255 1.320 1.388 1.458 1.532 1.606

1.000 1.032 1.062 1.091 1.120 1.149 1.178 1.208 1.238 1.268

1.0000 .9989 .9928 .9794 .9582 .9298 .8952 .8557 .8127 .7674

.5283 .5837 .6286 .6652 .6953 .7202 .7411 .7588 .7728 .7867

.5283 .5831 .6241 .6514 .6662 .6697 .6635 .6493 .6289 .6037

1.000 1.081 1.158 1.231 1.300 1.365 1.425 1.482 1.536 1.586

.9l29 .9870 1.057 1.124 1.187 1.246 1.301 1.353 1.402 1.448

.4082 .44l4 .4729 .5026 .5307 .5571 .5819 .6052 .6271 .6475

2.00 2.10 2.20 2.30 2.40 2.50 2.60 2.70 2.80 2.90

.5773 .5613 .5471 .5344 .5231 .5130 .5039 .4956 .4882 .4814

.1278 .1094 .09352 .07997 .06840 .05853 .05012 .04295 .03685 .03165

4.500 4.978 5.480 6.005 6.553 7.125 7.720 8.338 8.980 9.645

2.667 2.812 2.951 3.085 3.212 3.333 3.449 3.559 3.664 3.763

1.688 1.770 1.857 1.947 2.040 2.138 2.238 2.343 2.451 2.563

1.299 1.331 1.363 1.395 1.428 1.462 1.496 1.531 1.566 1.601

.7209 .6742 .6281 .5833 .5401 .4990 .4601 .4236 .3895 .3577

.7978 .8075 .8l59 .8233 .8299 .8357 .8408 .8455 .8496 .8534

.5751 .5444 .5l25 .4802 .4482 .4170 .3869 .3581 .3309 .3053

1.633 1.677 1.7l8 1.756 1.792 1.826 1.857 1.887 1.914 1.940

1.491 1.531 1.568 1.603 1.636 1.667 1.695 1.722 1.747 1.771

.6667 .6846 .7013 .7170 .7317 .7454 .7582 .7702 .7811 .7919

3.00 3.10 3.20 3.30 3.40

.4752 .4695 .4643 .4596 .4552

.02722 .02345 .02023 .01748 .01512

10.33 11.05 11.78 12.54 13.32

3.857 3.947 4.031 4.112 4.188

2.679 2.799 2.922 3.049 3.180

1.637 1.673 1.709 1.746 1.783

.3283 .3012 .2762 .2533 .2322

.6568 .8598 .8626 .8652 .8675

.2813 .2590 .2383 .2191 .2015

1.964 1.987 2.008 2.028 2.047

1.793 1.814 1.833 1.851 1.868

.8197 .8279 .8355

3.5

.4512

.01311

14.13

4.261

3.315

1.821

.2129

.8697

.1852

2.064

1.884

.8427

3.6

.4474

.01138

14.95

4.330

3.454

1.858

.1953

.8716

.1702

2.081

1.899

.8495

3.7

.4439 9.903 3 xlO.4407 8.629 3 xlO.4377 7.532 3 xlO.4350 6.586 xl0-3 .4152 1.890 3 xlO.4042 6.334-4 x10 .3974 2.416-4 xl0 .3929 1.024 4 xlO.3898 4.739_ xlO 5 .3876 2. 356-5 xlO .3823 1.515-6 xlO .3804 2.091 7 xlO.3781 2.790 12 xlO0 .3780

15.80

4.395

3.596

1.896

.1792

.8734

.1565

2.096

1.914

.8558

16.68

4.457

3.743

1.935

.1645

.8751

.1.U39

2.111

1.927

.8619

17.58

4.516

3.893

1.973

.1510

.8767

.1324

2.125

1.940

.8675

18.50

4.571

4.047

2.012

.1388

.8781

.1218

2.138

1.952

.8729

29.00

5.000

5.800

2.408

.06172

.8881

.05481

2.236

2.041

.9129

41.83

5.268

7.941

2.818

.02965

.8936

.02$0

2.295

2.095

.9370

57.00

5.444

10.47

3.236

.01535

.8969

.01377

2.333

2.130

.9526

74.50

5.565

13.39

3.659

.8990

2.154

.9631

5.651

16.69

4.086

2.3Tf

2.170

.9705

116.5

5.714

20.39

4.5l5

2.390

2.182

.9759

262.3

5.870

44.69

6.685

2.423

2.212

.9891

466.5

5.926

78.72

8.873

2.434

2.222

.9938

11,666.5

5.997 1945.11

7.631 3 xlO.9005 4.470_3 xlO .9016 2.745 3 xlO.9041 3.974-4 xlO .9050 9.753 5 xlO.9Q6l 3.255 8 xlO0 .9061

2.359

94.33

8.488 3 xlO4.964 3 xlO3.045 3 xlO4.395-4 xlO 1.078-4 xlO 3.593_8 xlO

2.449

2.236

.9998

2.449

2.236

1.0000

3.8 3.9 4.0 5.0 6.0 7.0 8 9 10 15 20

100 CtO

6

44.11

0

.8018

.mo

A2a:l

AIRFOIL DESIGNATION SYSl!lIS NACA WR 1-560 Systematic Investigation of Airfoils. Pressure distribution measurements over the upper and lower surfaces of airfoils led to the concept of investigating independently the variation of upper and lower surfaces, but this procedure was found unsatisfactory because modifications of either upper or lower surface 1I8!'e found to alter the nO'll' charaeteristics over the entire airfoil.

For example:

NAC! 65(lO)-211.

If the design lift coefficient in tenths or the airfoil thickness in percent of chord are not whole integers, the numbers giving these quantities are usually enclosed in parentheses as in: NACA 6$(318)-0(1.5)(16.5). Some experimental airfoils are designatea bY the insertion of the letter x immediately preceeding the hyphen.

Studies in the hydrodynamic theory of airfoils led to a concept of systematic investigatiem of airfoils based on 1leaning of the !lAC! l-series, the NACA 7-series, the changing first the shape of the median line and then the NACA Supersonic, the NACA Helicopter series, and the German thickness distribution about the median line. A good thick- DV1 airfoil designation systems are sh01lIl in Figures A2a-9, ness distributicn determined by tests em symnetrical airfoils A2a-1.3, A2a-lh, A2a-l$, and A2a-16 respectively. at 10'11' speed is that shown in Fir, • .l2a-l and this thickness OA __ -- ____ _ _ _ distribution was adopted in the testing of two large familie! ~ ~ _ -~ _ J( of airfoils knOllIl as the four-die it and five-digit series. -O'It----.:-:; - ~:'_-l

For the four dig! t series, a mean camber line con0 .2 .Ij .iI .e II) sisting of two parabolas, as shown in Fig. }_2a-2 was used. y. O.29697X- .126Ox - .3$1~ + .2483x3 -.101$x4 A typical airfoil of the NACA four-digit series is drawn to scale in Fig. A2a-3. This is the NACA 241$ airfoil. A diaFig. A2a-l HAC! basic thickness distribution. gram showing the meaning of the four digits used to designate f,r~t H-) 5e-cQ,'d the airfoil is presented in Fig. A2a-4. The meaning of the f'o.ra"o/F"".b.l~ digits of the five-digit series of airfoils is shown in Fig. A2a-$ and further explained in Table A2a-l. The mean lines of the f1 ve-digi t series are of two types shown in Fig. A2a-6 those with a straight trailing edge having a third digit of t.,. t.l'1o,.. Co",bQ' ,."fol'l'IOI'l zero and those with a renex trailing ed;e having a third digi t of one. Fig. A2a-2 liean camber shape for four-digit series of airfoils. 1lodifications of both these four-digit and five-digit series involving changed nose radius ard crAnged location of rM",d/Cin lin" the point of maximum thickness are deSignated by two supplementary digits preceded ~ a dash, the first digit after the dash designating the leading edge radius and the second arter the dash deSignating the maximum thickness location acconting to the scheme shoilIl iii Table 12a-2. The dash num6--iD Ii" -- '(,0 - 80 - -iito ber -63 is usually not written because it represents the normal leading edge radius and thickness loactions shown in Fig. !2a-3 Scale drawing of NAC! 241$ airfoil. Fig. A2a-l. For example 241$-63 is the same as 2415. Jlodified 0009 airfoils are shown in Figs. A2a-7 and A2a-8. 2 4 ~ median 'f.-........... )(ax thickness The dash number -4$ was found to correspood closely to per cent chord ~:a:rcent of the special shapes developed for high speed. 'nlus the 16-009 airfoil may be considered identical with the 0009-4$. Position of max median camber For ordinates of above airfoils see Teclmical Aeroin tenths of chord dynamics, 2nd. Edition. Fig. A2a-4 Diagram showing meaning of digits of four'!he NACA 6-series airfoile are usually designated by a digit airfoil series. six-digit number together with a statement showinl' the type of mean line used. See Fig. A2a-10. When the mean line used 2 30 1$ is obtained by combining more than one mean line, the design lift coefficient used in the designation is the algebraic thickness sum of the design lift coefficients of the mean lines used, in percent of as shown in Fig. A2a-ll. chord 1lean line Airfoils having a thickness distribution obtained by shape linearly increasing or decreasing the ordinate! of one of designation the originally derived thicknes! di!tribution! are designated as in Fig. A2a-12. Fig. A2a-S Diagram !howing meaning of digits of fivedigit airfoil series. '!he more recent NACA 6-series airfoils are derived as members of a family of airfoilS having the minimum pressure at the same chordwise position as the earlier individually derived airfoils, and are distinguished from them by writing the number indicating the 10'11' drag range as a subscriptJ for example, NACA 6$3-218, a.O.$. 'n1e designations of airfoil sections having a thickness ratio less than 0.12 of the 10 1$ 20 $ 2S chord are nO'll' given without the subscript number indicating the low-drag range. As an example, an !lAC! 6-eeries airfoil having a thickness ratio of 0.10 of the chord would be designated HACA 6$-110. If the ordinates of the basic • • thickness distribution have been changed br a factor, the 2.3 2.8 3.1 10'll'-drag range and thickness ratio of the original thicknen 3.l 3.7 4.2 distribution are encloaed in parentheses as follO'll'SI 6••••••••••••••••••••••••••••••• 4.6 5.5 6.2 NAC! 6$(318)-217, a.OS. If, hml9ver, the ordinates of a basic tliickiless distribution having a thicknees ratio less than 0.12 of the chord have been changed by a factor, the number indicating the 10'll'-drag range is el1ldnated and orU7 the original thickness ratio i. enclosed in parentheses.

;Ur;

1i::l~

~.9851

·993 100.000

! L.X. rao;Utl.!: 0.2;6 I Slope of raditl.! through L . .8.:

~1·935

~1·951

:U~ :i:tJ~

-1.z60 -1.0Z0 ··768

::Ht, ·090

I

.159

0 0.064

64-209

,-Station. snd ordinate, given in poercent o! &1rfo11 Chord]

percent ot a.trfo1l ebol"d]

Dpper

~urtae.

I

Lo.... r Surtace

64-210

Ordinate

:~ :1'9t -'ili ~:~i 1.719 5.066 ::~l7 7:~~ I t;.u 7.568 ,z.lJl' O 10.067 :i:~ib -1. 24 .')7' :~?l i

NACA

NACA 642-415 [Stationll IUld ordin.tea glTen in

NACA

r...ower Surface

Ordinate

-1.675

-6.129

64-208

&r:d ord:'ns.te~ gJ.vep in percent 0: airCoil chord...l

Upper Surface

~~:g~

45.02;

NACA

~3t"ticnl!

-4'8/9

-6.452

50.000

64-206

679 039 t·.011 2 25-0 4.066 '0.0199 ~~l' , ·991 45.009 "Oill 50.000 50.000 ~:t+o "'.91 2 ]5. 003 14.9 5

-6.345 -6.40.2

)5.

40.

NACA

i

i

rStatione illld ordinate, given in. - :>Je1'Cent or alrtoll cho~

I

Upper Sur!ac.

_~

~orti1nate

Statioll IOrd1.r::lata i Statlon 'Ordlnah '

I.o~r

Stat10n

Surface

ord.inate

Reference: Abbott, Ira H_, von Doenhoff, Albert E.and Stivers Louis S. Jr. Summary of AirfOii ~ta. NACA iJartime Report 1-560. , L.B. rediu.a: 0.579 : Slope of radiu.a throU&h L.X. t ~"TfONAL

.&DV!SOCIY

COMMITTU ft1 WOIIAIITICS

O.ct34

L.E. radilUl

0.7Z0

I Slope of radi\!.! through L.X, I

0.084

~ATlO"'.L .lDVISOAY COMMITTEE Fot A£lOIIAUTfCS

A2a:ll

NACA 643 -018

NACA 64 3 -218

~tatlcns IJ"lQ ordinates g1ven in

Upper Surrace

~tationl! and ordinatell g1ft..? in percent of airfo11 ehordJ

percent ot urtoll ehord-': Upper Surt&.ce

Lower Surt ace

Sta.t1on ,Ord.1n .. te I st.-;;i;;-fOl"I!inat;"j

Lower Surr.c,

,

Upper Surface

Ord.1tle.h : Station I Ord!nate

T

o

NACA 65,3-018

[Statl=1! and orCin.tea g1n_1'l in

percent of airroll chor!!

0

·50 ·75

1.25

0

:~~

'·5

.88,

Ug( •

1.~Ol I

2:,Jl ' 1.703 :

1

7·297 '

,,\:j&l. 9· 5 iI:~!1.• !:~ 9.760 ;Z.

;0

,,:~

'~66

{m

~6:gj~

II

~.814

~6:g~

1,.5. 0 28

50.000

klr

0

:2:0~

-1.~7

-·teE -. '

100.000 I

0

i ~~er~~J1u!·~:0\lg1l L.!.I

"------------------------

o.~

I

0

:~M

.95 0

2.152

4.609 7.095

i

b"og

d2~

~.657

~:m .10.176

jl:l:l

' 10·730 11. 0 37

1.Uil ·707

!

2·'70 ,·l57

Lowoer Surface

0 ·737 1.ol.4

U~~

5·391

16:lg~

15.lS3 20.343

25·29'

30.237 5·177

U56 ;·390 2.~ 1. • 92 : 0

70

~

I 1

!

i -;.;9 0 I

-2'ill

85 90 95

!· -1. · -. 92

100

0

1.92

L.E. rad1ua:

I

tipper S1.l1't ..ce

Lower S1.l1'race

Station 10rdinate

Statior., Ordinate

a-O.5

NACA 65,3- 618

[Station! 1Llld. ordinate! gi,.e~n in percent of .. irfo11 cb.ord~

Ordinate: Station ,Orc:l.1oate

I : !

i:Ul 6 ·• :l:i' :4.4~ .456

NACA 65(216)-415

!Stations a.nd ordinate! given in percent o!' airtoil chord]

Station

~.916

45

U~

1:~l4

i .8.990 _8:~~

'6

8.990

8.916

i6 85 90 95 ~

NACA 643-418 ~rsurtace

1

56

1:14\ 1650 65

16 65 70

-1· \' I

I -z.i2

2·5 5·0

dH

46 45 50

-7·535

i

lb· 5 I 1t:~ 15 . 20

i.

~6

-(:i;1

·9 § I < , ~.9 .9 81 :

':m I

100.000 I

20

:

I:i:m -2.

:15

!:~l4

15

!

lordinate ,0

1.25

2·i28 ,. 31

2·5

1..50

:1:11 :j:!lll

o

~:6il

I 10

:,.~c

·9t;

! 6 . 935 4

:15

:t:~

Station

0 1.!Z4

1.25

w7.011

~.l76

1

!

95·019 i

166

3P>§S I '0.0>8

l:!!j ! 2.62;

i6::;t

85 90

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0

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i:m 2:m 49 '1. 49 ./,6

1.099

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0

[;tation. and. orcltlll.te! ghon in percent or airfoil chord)

[StatlollJl and ordln.at .. glvep. in percaDt of airfoll cho:-d.... Upper Surhee

Station

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sur~

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Upper SUrface

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1.331

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65 1 -012

NACA 651-212

[at';;~:ll~~ O~;tJi· ora~:

~UppO'

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.....,

....1>0

\\ \\.

o o

~

.... ....

t::t

~ I'),-_~ --d, , ~( >
9.0 x '06.0 o ,.0 0- r--\: ~ Ill. Standard roughness

= 0.4

Ouej:lne~ s

:;! 1. 2

leP

6~-series.

I I

1.6

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6

1-1

NACA

R

,.-1

1.6

I

(e) 2.0

Ct

I

and 0.6

x

Airfo1l thickness, percent of chord

NACA four- and five-digit aeries. I

6

1 l:,

"".

x 106

~8St andard

~ t.---

G

""

12

R

9. o

i'--- hc:::::0 ,. o

r-o- h

44-seriea (4 digit)

x lOb

~Db. o

1. 2 1--T--t---;--+-W'-±~-+---I--+-*,,:J.::0

'--

16 20 Airfoil thickness, percent of chord

.-i

6

-(;)

t--+-:--+-=--;:;+---r-t--~;:"f'~-~~::A-(>

j:::- tQ:

x 106

o lr-l' o •o

2.0

I

R

t--

o andard V .>---- r-o-1'2-::::: g..L ~81lt ?

1.2 NACA 2;0-series

I

and 0.6

1.2 1.6

I

~

1

1.6

"""" ~ 1:::--.....

-........

I

= 0.4

6 x 100

I

100' 1.6

.8

R

L--'----"--

1.6~~--~~-~~-.--.---r

c

R

. 2.0

1.2

'--T---r--~-4~~-+-~~-i=G0,.0 b.O

x

=0

t1

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r-

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roughne~s

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.&if""

1.2

106

lOb

~ y+

o

(f)

(> 9.0 x 106

Ob.O

1--'---T--r-+-'-4l--+--t-+-~>--~'0

;.0

.-

~~t andard roughness 6

I

x 106

16 20 8 12 Airfoil thickness, percent of chord

1.6'--~-'--'--'--'-~-.'--~-'--'--r-'

1. 2

C-. ~

~

x 106

4

L-~~_ _~~

R

""'" 9· o ~. b. o t-. o

NACA 66-series.

l.& __------~------_.------_.------__, l.4~~~~~~------~-------_+--------_1

.£r-t---.~_ 1_.8' St andard 1.2 I ...!;;.,._--~~~~~~~...:::::""""'::::--t....-------I rou&me§s ~

r--+---r--+-~~~--~~~-+--1---r--

6

x lOb

.8 L--L-~4--~~8--L--IL2~L-~1-6~--~20--L-~24 0 Airfoil thiCkness, percent of chord (d)

NACh 64-series.

0.&

~

______

~

______

__ M-

-L~~

0.15 1.0 Figure 41.- Variation of maximum section lift coeffiCient with airfoil thickness ratio st several Reynolds numbers for a number of NACA airfoil sections of different cambers.

Reference: Abbott, Ira H., Ton Doenhoff, Albert E. and Stivers, Louis S. Jr. SUmmary of Airfoil Data. NAC! Wartime Report L-560.

______ 1.5 2.0 Dda on effect ot Mali! number on C.Lmax from AJ'H! 5773 lig. I.

~

~

A2a:20 IlINIIltJll DRAG OF AIRFOilS, Re • 6

It

106

NACA lIR 1.-560

0 NACA

o NACA NACA !; NACA VI NACA

008

o

>----I--, 004

r-- t---.. .

p osition

.6

.4 ·5 of minimum pressure, x/c

NATIONAL ADVISORY

·7

00

230

"il I

Rough

C>

Snooth

1.0

.012

~ (4- di g it)

0

.012 I - - -t:.

.2 .6 .8 .4 Design section 11ft coefficient,

c'1 Figure 11.- Variation of section minimum dr~g coefficient with camber for several NACA bseries airfoil sectiong of 18- percent th1ckness ratio. R. 6 x 10 •

OO}

0 c------ - 8

"

65.3-818

COMMI~TEE fOIl AERONAUTICS

Series

---r-

11s:: .0oB

"1

.004 :

Figure 9.- Variation otminimum drag coefficient with position of mini~ pressure for some NACA 6-series airfoils of the same camber and thickness. R, 6 x 10 6• .016

o 6 -

6~t:. 6 -

NATIONAL ADVISORY COMMITTEIE fOIl '1RONAUTICS

.,

0

NAGA airtoil o 6~-series

632-215 .012 642-215 652-215 662-215 67,1-215 .0oS

(5- di g it)

;>--'1 .0-

.)--'"

.r-' .rO-! :r-

--'7

.008

~oush

~

.004

Smooth

~~

~~

-

~.

'¢

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--

t--";

t

..to.

-----

°1 i

00 (I O. 1 o. 2 ~ £:'0. ~ "il O.

;.0-

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1

... 1

---

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(e)

NACA 66- series.

-

52.-

r

~

.26 Cl0

1

..--,

= 0.4 and 0.6

.128 0

.26

,r-

-

"-

I--

t '"- ~ l--

cOl = 0.2

.24

.28

1

1

0

.26

r

~

l~

C· 1 4

=0

8

12 16 20 Airfol1 thickness, percent of chord (d)

NACA

~

8 12 16 20 4 Airfoil thickness, percent of Chord Figure

series •

•28

.24

0

,ir"'"'

c· 1 = 0

=0

64-

NACA

"o0. ..."" .8 .28

8 12 16 20 4 Airfoll thickness, percent of chord (c)

.24

o

6

.2 6

= 0.4

] .28

....o

.,

1

Q

H

......,o ...

-

.,c:

~ r-

c· 1 =0.12

.24

c.

.24

...,"

-0

..., Q

v

Q

65- aerle ••

Reference: Abbott, Ira H•• von Doenhoff. Albert E. and Stivers Louis S. Jr. Summary of AirfoU Data. NACA wartime Report 1.-560.

Concluded.

A2a:25 VAXDl1J)( LIF'l' WI'IH FLAPS, He. 6 x 10

6

NACA lIR 1-560

Hinge location 1

2.8

~

~I 2.0

··

""g ~

1.2

~

: J

~

?

E 1.6

·

r-=

" IACA .612151-216 -7 V

2.4

V IL ;--- lIAC.

65,~-61a

, J!

.17. -----,

t (a)

I-

Pisure 55.-

Flap eontlgurat1oh.

Flap conf1surat1o n and llIax!m1.m 11t't eoefficienh for the Nne" 6;,1...420 airfo1l dth • O.25-airfoll-chord hingec alotted flap; R, 6 '" 106,

3. 2

Y '7

;/ I[)""

'I

/ ~.~

I

2. 8

.8

•~

f

2 •4

•• I-

j

NATlONo\l ADVISORY lTlCS

jcoooo'l" i """'t

.... 0

20 ~ Flap deflection, Or. dog

0

a:.

60

~t(;15~~i6~;~h:'1i~;~e~i~hrg~25~a1~~tl~~£~~1~1:n": "Ifo:6 xlrl-.

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I R

V: ,

/

.p.o

/

VII IIII; I'; I

II , ,

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;1// V

-/ ~f0

....eA 66(215)-216 JU.CA 2~012 It rom 2'eterC)Oe 5'7}

V

1.6

(approx. )

f-- KAeA ~ /-RACJ.

,./

6 x 106

~

""-

Flap-hinge location CD G

1

--

-

2

.V I V

~.5

.016

~

-

).

\

(\

8

t)

()

..... ...,

.020

t)

0

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\~

()

1. 0

)

~

Q)

II)

-~

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/

~

V

~

17 '-......

ru::..

NATIONAL ADVISORY'

I I I I I

Q

.0;

.04

v;co

where Q • flow through suc-

~

..... tion slot, fV/sec. .... ~

....

Vo. free stream velocity, rt/sec.

~

.008

'oo"

-.4

o

.4

6

6 .0 x 10;

~

~

-0.;,.

0 -1.2

-.8

-.4

y

~

:)

0

o

I

CQ.

-

0

.015 .025_

~ -a- -"'-'~

""

Section lift coefficient,

Figure 6.-

~

~

I

/

~ v

b • span, ft.

cL

model with standard roughness.

~

~

1.a

.8

Section lift coefficient,

~~

c • chord., ft.

i

I

o

~

...,-

_

COMMITTEE (OA AERONAUTIiS

Flow coefficient, Oil. -.8 Figure 17.- Effect of Reynolds number and leading-edge roughness on variation of maxL~um section lift coefficient with flow coefficient for NhCA 641A212 airfoil section with leadinged~e slat and double slotted flap. Os' 22.0°; x s , 0.0;6c; Ys' o. 037c " 0v' 16.5°; X\., o. 004c; Yv ' o. 014c ; Of' 55.0°; R x f ' 0.oL,4c; Yf' 0.005c; test, TDT 990. Jl .012 ~.

d

[7

.....t:T

-v

NATIONAL ADVISORY , _ COMMITTEE FOA AERONAUTICS

.02

/ p8

JJ.---'

.004

.01

I

...r.----€

5

a a

1/

.8

1.2

1.6

cL

R = 6.0 x 10 6 ; model in smooth conditiDn. Drag characteristics of NACA 64lA2l2 airfoil section with boundary-layer control. Tests, TDT 953, 984.

Referenoe: Quinn, John H. Jr. Tests of the NACA 641 A212 Airfoil Seotion With A Slat, A Double Slotted Flap, and Boundary-Layer Control b,y Suction. NACA Teohnioal Note No. 1293.

A2a:28 EFFECT OF SURFACE IRREGULARImS ON WING DRAG FOR COllPLElE FllLL SCALE AIRPLANES

lIR 1-489 .00.3

:r.A13LE IV.- Wing Profile Dr"€s CDo

Doscri!)tion

!:lDC.S-

urcd

CDo smooth 1'IiIl€; (cst. )

lIGno

1

-. -x_

I

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4' 10

jmil/i~

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,

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:

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l

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number, R

i I

!

I,'

~Gt~l

11 "cod, filled

5

. -3

6 rowe of 3/32" brasier head rivet. on each .urface of 5-foot chord airfoil Pitch 3/4". forward ron, 52 percent of the chord from leading ed&e. 2 13 rowe of 3/32" oounteraunk riveh 011 each .=face of 5-foot chord airfoil Pitch 3/4". lonard ron, 4 percent of the chord from lea.din& ed&e. 3 e ron 011 tap and 6 rowa 011 bottom lJ'.ll'face of 5-foot chord airfoil. Pitch 3/4". lorward ron, 36 an:! 52 percent of the chord from leading ed&e 4. 6 Joggled lape facing aft on each surface of 5-foot chord airfoil. Forward laps, 8 percent of the chord from leading edge.

.0041

.0060

+

]X>1nt.

+~

1

0

.0079

4

RrtynO/d3

,,

, I'"

, ' ! I: !I -]-1-- 3'- i~i~::-]-I-- -3I.:;,

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1,

-.u l~ cttl1tt"'~'\.uj~,

l.&rgitl"""

tMt bol4.I ~t..lJ'

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·'~J"I~.t~;

--.t

oorw ~

NATIONAl. ADV'JSOQY

aill!'t-e.1'Q3 2_182

_426 .430 .426 .)74 ·Jl5 ·272

·409

7·673 7 ·7"" 7·664 7-353 6.728 5-661 4·894

·[U;9

.746

.828

·092

6.615

·562

·562

2-490

2.140

3·310

2.)20

3·625 4.170 4.431 4·512 4·550

2·390 2-'75 2-532 2·578 2_600 2-57" 2.470 2.260 1·902 1.64"

....

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1/,.

I

..

I\.-facA

(o)aC,-O.67°

~

•

~

I

_

~

;

~

number., 1'4.

Variation of windshield drag coefficient with Mach number.

x-t. X-I

----~

-.

r f'A

I

L I

>..c""

2

x

-

1- .-----'-t-~. _'t -1 r-'~ _ L

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f

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I

1- {~;~021

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17{ e»

(7{b»

__ + ___+__---+_______

0.0021

installed In rear underslung fuselage duct

'r'1~-rolj-8x.19

t:::I

~

0.0005

0.0005

17{ a»

OIL-cooler lnstallatlon

_____ ~_

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0

0.0004

--

Carburetor ",oop.

.0284

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(6)

Coolant-radia.tor lost_ll.t1cn

i:

12

0.0041

(5)

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I

I

1)

O

Wing-duct inlets

--------

I Vanes

10

0:021~-=¥~10

(..0219

0.0163

.0264

AC

Cowling-flop and hlnge-llne-ga.p leakage

,

EG

7

~

AirplAne in s .... lc. condition

ij ~o

5

Alrplane sealed and ialred

i1:x: rf ~ ~lJ>

r·-I4

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0.0040 (21{a»

1

I

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1

I

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LandIng gear Annament

(27(cl~+

----------------~----~-

1

,.

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Tall .... urface-g.p leakas.

"O{a»

Tall wheel :'nd _rre.tIng hook Canopy modl!"lcatlon Rad10 antenna

e»

(}9{ .0004

I

.0004

---+-·,--1---+----

(}9{ .. ) )

-Estimated drag coetfic1ent ..

bModltlca tlons.

> .,0-

I\.)

\no

A2b:6 AIRPLANES TESTED FOR DRAG

NAeA 'lIE 1-108

(h)

AirplRn~

8. Sealed and fa ired condition; prolceller removed. _ _..;;.;-" area = 442.0 sq.

Service condition; propeller removed. Wing area" 490.0 sq. ft.

Figure 1.- Airplanes mounted for tests in Langley f'ull(1) Airplane 12. Sealed and faired condition; scale tunnel. propeller removed. Reference: Lange, Roy H. P. SUmnary of Drag Results from Recent Langley Wing area = 2113.2 sq. ft. Full-5aale-Tunnel Tests of ArlIlY and Navy Airplanes. NAGA wartime Report

1-108.

A2b:7 COOLING DRAG DATA NACA WR 1-108

"CD. 0.0040 flV. l7 ..ph

Q, 16. OOC au tt per .. In H - Po. 0.4Oqo

(a) Orlginal long-nos. cow11ng. Flgure 3. - Cowllng drag on airplane 1.

II

----r----':::::2-1---: ~ -~~" Flgure 10.- 11ng-4uct inlet on airplane ll.

Od-cooler oui:lei

'~i

dCD. 0.0021 flY. 12 mpr

Original coolant-duct inlet

. ~~.A .....~ (' ~",J '\' 4CD.

dY.

o.ooi

14

mph

. ;,.

.' 'IIIIiIII ~ .... /

Revised coolant-duct inlet

,.

\I

.

\

.,-

~~~/

,/

/'

Q/Vo ' 0.12 H - Po. 0.40...,

)

(al Orlglnal duct lnlet.

.~;.

7

(a) Coolant ducts.

flV. 5 mph

Q/Vo • 0.13 H - Po. 0.95

fJ t:I

~

t,""".l

., t:I ~

>

WiR(,

Ballii'

as

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T

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-----

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1. RI

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Poor :t~r('('nlf'nt O\'l'r whol 0 ...;J cf- 00 1-'- '1

306 Q

i .• = _4°; othrr-I 0

~

'1

0

t:I ) For mud guards, adti~or each wheel, not per sq. ft., 5 to 10 8. Engines (Radial) and Nacelles (per sq.rt. of circumscribed circle) (a) 240 HP 7-cylinder (J-5) engine, cylinders half exposed, no cowl 14 to 16 (b) Same as (a) but with N.A.C.A. Cowl 3.4 to 3h (c) N.A.C.A. Cowled Engine, including interference with wing. Engine well above wign, exposed struts. 9.5 to 14 Engine well below wing, exposed wing. 5 to 7 Engine crankshaft on wing chord 2.0 to 2.5 7.

)

8.1 OALCIT data on Nacelles, Fuselages, etc. (JAS Dec. 1936) Drag added by Fuselage (based on frontal area) Bare hull of dirigible "Akron fl outline, circular sect. 1.8 Large camnercial transports, low wing, no nose engine 1.8 - 2.7 LArge military airplanes, no nose engine 2.3 Low wing, single engined transport 2.6 Low wing. small airplane with cockpit 3.3 enclosure Drag added by Nacelles (includes interference) Nacelle mounted externally above wing 6.4 I.eadi1l& edge nacelle, small airplane, relatively large nacelle 3.1 Leading edge nacelle, large airplane, relatively small nacelle 2.1 Drag added by Tail Surfaces .22 - .31 Single engined, low wing Multi-engined, low wing .15 - .28 .31 - .46 High wing monoplane or biplane

9.

Radiators, per sq.ft. Without shell, 5" to 9" deep Without shell With streamlined eow1

17 to 19 26 8 to 12

10. Struts and Wires (per foot of length) Round Hard Wire (Estimated from various sources) Diameter

.0641 .0808 .1019 .1285 .1443 .1620 .1819 .2043

B.S.Gage

14 12

10 8

7 6 5 4

Drag per ft. .12 .15 .21 .27 .31 .35 .39 .44-

Two Fittings .4 .4 .5 .6 .7 .8 1.0 1.2

Stranded Aircraft

C~b1e

Diameter Actual

Nom.

1/16 5/64 3/32 1/8· 5/32 3/16 7/32 1/4 5/16 3/8 1/2

.062 .078 .094 .125 .156 .187 .218

.250 .312 .375 .500

.A2b:ll

(From Diehl)

strength #Ultjmate

Dra.g per. ft.

480 550 920 1350 2600 3200 4600 5800 9200

.16 .20 .24 .32 .40 .48 .56 .64 .80 .96 1.28

Two fittings

.42 .42 .52 .65 .82 1.01

2

1.21.

1.50 2.05 2.60 3.70

10. Wires and Struts (per rt. of 1ength)(Cont'd) Round Steel Tubes(Estlmated from Incomplete NACA Data.) Diameter Nom. InChes

Drag Two Per Ft. Fittings(Est) 1 2.2 2 2 4.4 5 3 6.6 7 4 8.8 9 Streamline Wire (From Diehl) End Streamline Strength Drag Two Diameter Section Threads #Ultimate Per Ft. Fittings. (Est. ) .138 .Ob~.192 .30) 6-40 1000 .033 .190 .064,x.256 10-32 2100 .044.43)10 .58)to .056 .250 .087x.348 1/4-28 34.00 .3125 .llOx.440 5/16-24 6100 .76)15 .067 •077 .97)rt • .375 .135x.540 3/8-24 8000 .085 1.23)of .4375 .159x.636 7/16-20 11500 .500 .183x.732 1/2-20 15500 .092 1. 53)wire Streamline struh Navy Struts,3:1 fineness ratio

Two

Lbs./ft.

Fittings (est. ) Thicmess 1/2" .092 .7) Thickness 1" .138 1.1) .180 1.8) Thickness 1. 5" Thicmess 2" .220 2.2) Thickness 3" .295 3.5) Thickness 4" .36 4.2) (For built-up fairings, add 10 to 30% to above figs.) 11. !.liscellaneous Items,Lbs.at 100 MPH Control horns (per aq.ft. frontal) 8 to 15 (Add 100% interference if on upper surface) Fittings (air scoops, stops, hand holds, etc) 60 (per sq.ft. frontal area, including interference) Tail skid 2 to 5

a

Lbs./sq.rt. at 100 mph 10 ~ Hemispherical anemometer cup 36 ~D " II" 13 - . From NACA TN 450 at RN - 120,000

re

Cylinder drag data from NACA TN 3038 M
'

~

4

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3 fb/V

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II IICCOo~CC~@@8,·,',bPp'.

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11,0CKlIlOO\t.aoo,2.8.50~36'

' I 1108 2.50*11,225 I I

,I _~LC::·OC',',~~.H'@@651'5hP.p. il4oel125 h )li6 lSI

Pia boy BAJA'

I

11.57

li

Junior

SlNIak Aeto Corp. P. O. Boz £81

I~:f.

2ICon.C90@OOhp.

21Lye.0320@150hp.

Siybaby

1.18'300 20,000

I,

I,

ILyc.G0435-C2B@Z60bp.

,I

u"-

PA-18 Super Cub PA-1896SuperCub 'PA-I8-A 135

I 675160 Z,5.'iO i50: 80 !3,3.50 ,2,030 ,36'

P9,SCO

1 Lye. 029Q.D@130hP.\'I2.5iI151,700....7OOII4011950h300I,:H,'!l810'1150e$11,MO

ICon. CH5-2H@14:ihp. lCon.0470-E@225hp.

I,P.-\-20Pacer iPA-~2Tri-Pa.ce!'

StitsAircraft P. O. Bex MJ8'; RiKTlide, Calif.

~.;o' 1,460 .36 :6 6'

1690 :15,500......,

lCon.C145@145hp. 140 1120 : .. ICon.0470-A@225hp. 11115 1160 I 1..ltiO IJacohllRi55@300hp. ,\~ 165 1 I !1.200 2con.0410-B@2-whP'lm2051'111.7°O

170

Eahp.-Equivalentshafthonoapower.

HII.-HOI'a&I)(rWW.

lb.l- Pwnds thrust.

I

EstImatedd.atL

~

Former Chua Alnnft Co.altI!lane to be prtI(!uced

by FalrchHd. I

R1V-2 wHl have P&W T34 turbo9roPs.

$3,060 $8,950

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47G

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61 200

Ft. Wm-Ilt, Tf;t,.

Brantly Helloopl.. Corp. %l'ennEltulicCfI.

3V?~V~':;~2,~",:::"

8-130 XH8L·1

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Manufacturer and AddreSi

1

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700·1

,.

8001111

1~1'·"'f3:""I_"'I"'I·:""I':"".I:

I: 13"II'I,531 'wl ,...1"'\' 41

XV·3 1

7411118123'

300+121 I 1,2110

21 I Lyo.@IMhp.

B-2

1 ... I17'wI3'WI 6'IO'

PhllatUl"hia, Pa.

C. .NI AI""H Co. IIJiropltr f}if'iaiOI1 Widiita., Kan •.

CR·I

CI

~)

Convertawlnp. Inc. Amitvn"Ue, N. Y.

QlIadrotnr

Doman Hellcoplers. lilli, Dan/'Ilr". Conn.

L1r15

O",odyonfl Co. of America, Inc. St. James, 1,.1., N. Y.

2C 3 24

YR-31

UH-12R m·1

l'al"A,J/o, Calif. J~

'04'

Kaman Alrtn.ft Corp. WirnflarTJOrk" Con".

K-WJ

Alnnft Envln" Co. PoCUtotun, Pa.

K-3

Helidyno

H·2.1B Yn-32

MoDonneil AI""" Corp. St.IMlt., Mil.

Au_I Helloopier Corp.

MC'....C

."

Mtlrion,Pa .

•

PD·22

PH....' PD·lS PD·1S 8IJt:onky Alnnft Dfvlslon

B-1S2

United Alnnft Corp.

S-M

B~,CMln.

B-lI.

I ~ Con. COO @ SIS hp.

'II

1,0&0

4 I4

1 Lye. 80580-D @400hp.

104

18 3 73

P&:W IUlatl@400hp. 2 P.tW RI340 (it 1,200 hp. 2 Turboprop 2 Turboprop

HiO 1t3 304

1,300 1,300 ...... f1

'" 1""'°

1,0lI0 .... __ 1,000 •. 2,31!O1..

3110 1H IlOO

4,410

1

IBIS 12,800 833 70,000 .... 81,600

7,800 J 21,629 2 ....... 2

2 112' 2 110' ., 8S'

""1'001"8110 1"""'1'1'1'"

:::: 38'

all'

52

.. ' ....1·.7001

1,200

t30

40 Il,too

I ',180 121'

,, 1

311~inltlO2·2@1I1OhJl.

1"0 h.oool •. oool

200

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411 JI'r. flV8-2IS-BHIF fj)' MIS hp..

2

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.. 31' 5'14'2'19' 3'

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I

IH,8f7 1',110 jll.1

~

... 110' SB'

,.. 1".·' I'rT' 23' IS 221'

31

MO 13&

10' 15'

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770

830 4,000

711

A'

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.... 112'

151 tJaoobeR7M.ER@3lSOhp.

,""'_M.-"",-GO@>""hp_

YH-IS H.IOA XHR2S·t

48'

I" 1 1.....1'''' 1"1 ,.... 11.... 1'

HiII~

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4

~

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'21,Y,-&Vt-:IIlII-C3II@ ... ·,2 ramjet@ 32Ih.t. t Lye. OWl @ 340 bp. 1 PAW RI340@ 600 bp.

XHCH-t XV·II

.:30' &"130'8"'15' 0"

I la'

". I ...1'·0001 ____ ·_1"'1'·000 ,..... , , 1'1"" ~,860

3

HTK·I fiOK·l

K-' McCullocIt Motort Corp. Alnn" Dlvltlon Lo.AnoNeI', Calif.

2

, 11 Lye. S0580·D @ 400 hp.

alS' Hm.. HeiIOOfIf.... Inc.,

I

200

12&+

1__ .130.,..,+,.

111]82' 3

'..+1 _____ 1__ .___ 1""'1100 '·0lI0+1··8,600 _____ 1'I 100 300 1'1",000+

140+ I,m 15,800 100 103

1,000. ..... 1,000 .•....

11,100 6,100

3,"28 4,182

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O-tZ I p&:W Rt340-StHI@ 000. h~ 111 ZS· 2 P4:W R2600@ 1,900 hp

fOO· 1110 400· 1110

..".

•.•• 10'

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'.700

:177'

.·,.·.·1..'

77' II' 23'11' 25' 52' 11"14'2'16' "1S2' 11'14"'111'

32'

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GREAT BRITAIN

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TransllOCt

D.H.C.2 D.II.C.3

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______ I_____.I ___z_~ _~. ______ ~ __ ~_ ~___~__ AUSTRALIA

~

"TI

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Common.... 1h AIrMft c.,.

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a1110 rmkllli tnl Ihr....,au. Sub-VtB IInd..- a .Imllatoontrul

A2f:9 GENERAL LOGARITHMIC CHART OF POWER REQUIRED

•

1

..

•

1.0

Ii 6 .0

125 j....-o .....

15

~

25

20

115

Vv vi/VV1/[; ~l--vV- Vv ~V ~!-,j....-- V V- I----'Vv Vv VVVV ~l--!-- V I----' I----'Vv Vv VVVV V~ 1/11

3.0

6.0

Vv VV' VVl/V 11[; [J

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5 .s

I.0

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5.0

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Reference line for T~D/Q on chart of TD/Q vs QS for (3 • Const.

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lI;

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p. A2e:3 I! II

I

I I

I

ii

I

I!

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t\. ~ Reference line for fel~D2 on

I

1.2&

1

J

1

lo~arithmic

propeller e ficiency chart for Cp • const. p. A2e:6

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,

I ~

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V

30

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15 1.15

2.0

2.&

3.0

General lo~arithmic plot of power required for level fli~ht. Point ()l is the point of maximum LID ratio. To be superimposed on logarithmic propeller efficiency chart as instructed. Adapted from Boeing Aircraft Co. Chart by G.S. Schairer.

A2f:lO

PERFORMANCE WITH CONSTANT SPEED PROPELLER F~ctors

for calculation •

.9

>0

V

.8

°

q

.7

°

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/

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RESISTANCE, LB.

_-------_

36,000

-------

MAX. AMPLITUDE, DEG AFTERBODY KEEL HORIZONTAL

L

-b- = 5.2 24,000

~

--

TAKE _

STALL

8

I

I

6

I

I

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5.5

I

--

12,000

DEPTH OF STEP, "10 BEAM

I

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OFF~ 2

7.2

10~r------'~----.-----~----~~ 4 6 8 10 12 TRIM AT CONTACT, DEG. Figure 4.150 120 60 90 30 Typical effect of contact trim and depth of step on SPEED, F. P. S. amplitude of skipping during landing. Figure l. - Typical effect of increasing length-beam ratio with GROSS LOAD, constant length-beam product on take-off resistance. LB.

CDmin

.0075

Cma

LIGHT SPRAY

100

.0062 PROPELLERS CLEAR

80

.0067

.0051

.0012 60

.0060

.0044 .00135

R.N.'3.40xI0 6

40T--------r-------.-------. o 8 16 24

Figure 2.- Effect of increasing length-beam ratio on the aeroSPEED, FPS. dynamic characteristics of hulls as determined in the Langley Figure 5.7 - by 10 -foot 300 -mph wind tunnel. Typical effect of gross weight on the range of LOWER LIMIT UPPER LIMIT spe~ds fo~ spray in the propellers of a model of a twinMAX. PORPOISING PORPOISING engrne flymg boat. AMPLI TUDE, _ 'H I:l 6

l

:1

DEG.

\

.~~ ~

I

r

STt

ALLOWABLE_____

~ ~ACNGE·I_ LEM

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~--------_.__------+-.£-.---.....,....-----__,

24 POSITION

28

32

f/

_ _ _.

40

36

~

~

(.)

!;i a:

OF C. G., PERCENT M. A.C.

Figure 3. - Typical effect of longitudinal position of the center of gravity on amplitude of porpoising during take-off with constant elevator setting.

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Herrmann Engineering Co. Glenclale, CaJif.

Cam Engine

Jacobs Aircraft Er.~;n. Co. Po_,Po.

R765-A R765-B

Lycoming Diylsion of AYCO Mfg. Corp. StraJ.ford, Cr:m".

826C9HDE' 853C7BAl' 863C9HDl' 865C7BA1' 866C9HE1' 867C9HE1' 871C7BA1' 896C9HD1' 895C9HE1' 023.'>-Cl 0290-D2 0320 043.'>-A 0435-H 0435-Kll 043.'>-V G043.'>-C2 GS0435-B G0-480 GS0480-B GS0580-C 805SO-D' GS0580-D S0580-V

Pratt & Whitney Aircraft Diy. United Aircraft Corp. EIUt Hartford, Conn.

Wright Aeronautioal Diy. Curti .... Wright Corp, Wood.Ridue, N. J.

Bmep. G Ho R&d -

RZOOO-D5 R2000-2SDI3-G RZ800-CA3 R2SOO-CB3 R28OO-CB4 R2SOO-CA15 R2800-CAI8 R28OO-CB16 R28oo-CB17 R4360-CB2 745C18BA3 749C18BDI 826C9HD3, &; 5 826C9HD4 836CI8CAI 853C7BAI 863C9HDl 856TCISDA 1,2' 865C7BAI 866C9HEI 867C9HEI 871C7BAl l 956C18CAI 957C7BAI 968C9HEl 972TCI8DA1,2 975C18CBl 976C9HEI

Brake mean effective preasur .. Geared. Horizon tally opposed

SL

Radial. - Sea leyel

Vo

- Vertically opposed.

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4 4 6 6

Ho Ho Ho Ho

Direct Direct Direct Direct

902,31XJ, . loo2,5SOI.. ISO 2,600: 165 2,800:..

: 6 6 6

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2003,1001... 2003'100'1...... 2453,275....

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Direct Direct Direct Direct Direct Direct Direct Direct Direct Direct G .69:1

65 2,300 . 85'2,575i 952,655, 125 2, 5SOI· 1452,700 ISO 2,600 1 2052,600 225 2, 6SO 225 2,600 . 2402,200 260 3,100

1

1 ~~ C ~5 ~ --------1----------1-1- - - - -- --,- A1rcooled Motora, Inc.

~

4 4 4 6 6 4 6

6 6 7 6

~~

gir~~1

Rad Ho

i~;::1

R755-11

RI300-1A R1300-2

......

73 73 80/86

73

.. .

80/86

80/86 ........ .. ... ..... .. ' ..

80/86 SO/86 SO/86 SO/86 SO/86

1

9

R'd

9

Rod

8 8 14

14 18 18 18 18 18 18 18 28 18 18 9

9 18

7 9 18

7 9 9 7 18

7 9 18 18

9

Rad Rad Bad Rad

G .500:1

G .500:1 G .450:1 G .450:1 G .450:1 G .450:1 G .450:1 G .4SO:1 G .4SO:1 G .375:1

Rsd

Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad Rad

I I

1,45(112,700 1.2002,550 .[,4502,700 1,200 2,550 2,40012,800 1,800 2,600 2,40 0 2,8001,800 2,600 2.50),2,BOO I,S002.6O" 2,400IZ,800 1,60) 2,60" 2,400 2,800.\,60J 2,6Qn 2,400 2,8iXJ,1,5OO 2,60" 2,500 2,800 1,700.2.600 3,500 2,700 2,660!2,550

AVIATION WEEK,

15,1954

6.3 6.3

131 139 143 138 141 143.. ... 135.. ... 143.. .. 143 130...

170 182

......

...

144 132

04.95'

6.80 6.2 6.80 6.2 6.80 6.80 6.2 6.80 6.80

229 7.21' 187 7.21 229 7.21' 187 7.21 229 7.21 236 7.21 187 7.21 229 7.21' 236 7.21' 140 137 138 136 131 140 140 140 161 8.6 136 175 11.27 156 7.91 166 7.91 166 7 91 166 7.9

6.75 7.51 7.25 6.5

7 7.3 7.3 7.3 7.3 7.3 7.3 7.3 7.3

7.3 7.3

04.95' 04.95' 55.6' 60.45' 54.95' 56.6' 60 .45' 04 .95' 55.74' 50 .45' 55.6' 50.45'

04.95' 56.6' 56.6' 55.75'

~~~

265 263 355 385 352 485 431 264

1,3SO I,On5 1,350 1,056 1,404 1.445 1,080 1,385 1,460 236 260 268 392 362 383 396 422 488 432 492 592 58 597 585 1,585 1,605 2.317 2.3jj 2,357 2,350 2,350 2,390 2,390 3,682

49.10" 49.10' 52 80' 52.S0" 52.SO" 52.80' 52 80' 52.S0' 52 80' 55 00' 55.8' 55.6'

~~~

50 50'

6.0 6.0

55 74' 22.53x32' 22. 81x32. 24" 24.68x32.24' 29.59x32.24' 24.65x32.U' 28.21x33.12' 33.12' 23 .59x33 .12' 32.36,33.12' 33.12' 33.12' 30.75x33 IS' 30.75x33 IS' 30.75x33.18' 33.1S'

1

March

307 297 372

::~....

44' 44'

04.95'

1

'Or 10.14:1. • Or aS9:1.

152 152 140

7 7 7 7 6.1 7

50.45' 54.95' 55.74' 50.45'

.4375:12,200 ,2.800 1,40012,300 12, 100' I(f) '130 6.125x6.312 43751 2,500'2,800 1,470 2,400 11,000 100'130 6.125.6.312 ,666:1 1,42512,700, .. .. . . . . . 100,'130 6.125x6.875 .' 1001)30 6.125,6.875 56251 1,4252,7001 G .43751 2,700 2,900 11511451' 6.125x6.312 G .5625:1 80012,600" .... 91/98 6.125x6.312 Di.rec:.. 1,4252,700! .............. 1100/130 6.12~x6 875 G .4310:113,2SOI2,9OO: ................ 115/145, 6.120x6.312 Direct 1 8OO IZ,600i . . . . . . . . . . . . . 91/9816.125,6.312 G .5625:111,~7~:2,8OO! .... 1.. "'1 ..... !loo/13~1 6.125x68~~ G.5625:1 1,020 2,8OO, ......... , .... "1115/1401 6.125x6.810 Dlrect_ 800 12,600, ......... 1 ..... 91/98 1 6.125x63121 1 1 2,400 112,100 115/145, 6. 125x6. 3121 G .4370:12,700;2,900,1,600 1 G .5625:11 800[2,600 , 490,2,130 10, 500 [ 91/98 ! 6.125X6. 312 1 G .56:5.1 1,475;2,800, 890,2,400113,000,100/130, 6.12~x6.875 1 G .4315:1,3,250,2,900,1,910[2,400,11,400,115/145, 6.120x6.312 G .4375:1 2 ,8OO:.2,9OO 1,600 2,400.12,500,.115/145: 6.125x6.312 1 G .5625:1 ,535,2, 800 890 2,400,12,200,100/130 ,' 6. 125x6.875i I .1 ': 1 1

i lil

8.5 8.5 7.5

~.3

04.95'

6,40011001130 573.5.50 5,000 100/1301 5.75,5.50 6,000 100/130 575,6.00 8.5OOi1oo'130 5 75,5 00 8,500,108130 575,600 16,{)(X} 100.'1301 575',600 14,500 1OO'13015.75X6 00 16,000100/130 575.600 14,500 108 1135 5.75,600 5,5(f) 108i 135 5.75,6.00

1 for helicopter Installation. 'Manufactured under Curtiss-Wright lieense.

·Or8.67:1.

~

200 27 278

7

60.45'

G G G G

Rsd

Rad Rad Rad Rad

80 80

I

=

138 .... .. 138 ..... 136.... 140..

8.5

5.25,5 525x5

G .666:1 1,4252,700. .. .... 100/130 6.125.6.875 G .5625:1 800 2,600 . 91/98 6.125x6.312 Direct 1,4252,700 .............. 100/130 6.125x6.875 Direct 800 2,8001 .... ..... 91/98 6. 125x6.312 G.5625:11,4752,700i.. """ 100 / 130 6.125x6.875 . . . . . " .. 115/145 6.125x6.875 G .5625:1 1,5252'SOO'1 800 2,600 ............... 91/98 6. 125x6.312 Direct G.666:1 1,4252,700 ... " . . . . . . loo/1306.125x6.875 Direct 1,5252,8OOi ... , ........ " 115,'145 6.125x6.875 Direct 1152. ,81XJ1 812,350 6,500 80 4.375x3.875 1402,800 100 2,350 6,000 80/87 4.875,3.875 Direct Dmc\ 1502,7001112 2,4SO 7,000 80/87 5.125,3.875 Direct 1902,560 1422,300 6,500 80/87 4.875x387.; Direct 20G 2,800 1422,300 6,500 80/87 4.875x3.875 Dired 260 1,400 1871.200 9,000 91/96 4.875x3.875 260 3,400 1 1963,100 5,000 80/87 4.875x3.875 Direct 260 J,4(f)1 1802,7SO 6,000 80. 187 4.875x3.875 G .642:1 G .642:1 300 J'4(}~1 19; 2,760 10,000 91/98 4.875x3.875 G.642 2803,400 1952,7SO 6,000 80/R7 5. 125x3.875 G .642 3403,2001 210 2,75018,000110),'130 5. 125x3 875 G.642:1 375'1,300'1 2402,76013,500 91 198 487;;x3.875 Direct 400 3,3001 2R~ ~.750 1l,00D 100 / 130 4.875,3.875 G .642:1 4(f) 1.300, 232 2,75D 11,0001100.'130 4.875,3875 Direct 400 3,3001 280,3,000 12,000 \00/130 4.875x3.875

8

~

JS

~f";i i ~

7 7 7 7

~5

3.25x3.75

Rad

Rod Rsd Ho Ho Ho Ho Ho Ho Vo Ho Ho Ho Ho Ho Ho Ho Vo

... ...... ....... .. .. . ..

3 875,3.625.. 4.062x3.625 4.062,3.875 ........ 4. 062x3 . 625 .. .. .. .. . 4.062x3.875 .... 5x4 ..... .. .. 5x4 ....... 5x4 5x4 5.125x4.625 5x4

Rsd

6 6 6 8

R33W-34'

... :::::

4.5x3.5 4.5x3.5 4.7x4

7 9 7

6

RI820-76A,-76B R1820-101 R33SO-26WA RI3OO-1A R1820-103 R33W-30W R13OO-2 R1820·SO RI820-82 R1300-3

...

9

6 6

..

!~~;:~

91 91 80

...... ::::

Direct

6

0580-1 0580-3

..... .. ..

4.5x3.5 4.5x3.5 4.5,3.5 4. 5x3. 5

Direct

9 4 4 4 6 0435-4

..

80 80 80 80

R.d

9

c

g ~

- - ---=--1--=---1-----1-- - - - - -2-0

Rod

Rad

i

~~ c

co

7

Rsd

300 2,200 2752,200

!~

j

1

:,,'

].~~

i-g

I!

-!.E $ ~ ~ ~

7

7 Rsd

R13oo-3

,g.~

200 1,900

12 ......

I

I

" I

5.5 6.5 6.8 6.8 6.7 6.2 6.SO 6.70 6.2 6.8 6.8 6.2 6.7 6.2 6.8 1 , 6.7 ! 6.7 68

185 211 229 229 220 187 229 265 187 229 236 187 220 187 229 265 228 236

6.46' 6.46' 7.21' 7.213 6.46' 7.21 7.21' 6.46' 7.21 7.21 7.21 7.21 6.46'1 7.21 7.21 646'1 6.46' 7.21'1

1

• 856TC18DBI (R33W-30WA) is similar but has takeoff power (wet) of 3,500 hp. and dry gross weight of 3,445 lb. 868TC18DBI (R3350-85) Is similar to -30WA but weighs 3.438 lb. dry, Some have manifold absolute pressure regulator. 7 Military engine rated at 3,700 hp. max.

2,780 2,944 1,3SO 1,380 2,848 1,065 1,380 3,408 1,056 1,3SO 1,445 1,0BO 2 ,957 1,065 1,390 3 ,514 3,029 1 ,460

U.S. AND FOREIGN GAS TURBm~. l"ARCH.

A4.1e

1954

U. S. Gas Turbine Engines

AJIlson DIvision General Moten Corp..

IndiaMpolia, 1M.

Boeing A1rpjane Co. St4J;lU, WaM. Fairchild Englnt Olvidan,

Fa~~!~nt~I.~I~~y~ Corp.

GMara! Electric Co. AIrcraft: Gu Turb[ne Division Cincmnati 16, Olt.io

Pratt & WhltneyAlra-aftDlvlsion UnUItd Alrtn.ft Corp.

EallHl;lfard, Conn.

WeeUnghguse Electric Corp. Aviation Gat Turbine Div!slofl Phil.Gd&phi413, p~

.. -Appro:dmale. AFJ -Ax\l.l.fiow turbojet.

AFP -AXial-Row turboprop, CFJ - C8Illrlfupl..ftow turbojet.

E.shp.- Equiva.lent$haft hOl"MlpOwer. I.b.t. -Pounds thrust.

'At 70% cruill8 power_ \ UncowIed., minustallplP6 but with taileone.

Leading Foreign Jet Engines Oeslgnation

Manufacturer and Address

AUSTRAliA Commonwealth AIrcraft Corp. Pry.. Ltd. PDrlM!lbcvrM

CANADA

A. V. Roe Canada Ltd. Mallan,Ont.

~ene2-VH

A.vOQMk.l Orenda 8 Orenda 9 Orenda 10 Orenciall

Mamba ASM.3 Double Mamt. ASMD.l

5i.83-

".,. 1.,..

Pythoo Mk.3 SappblrtAS8a.6

-'5.,.

106'

ViperASV.3 Viper' !BV. 9 Brlml Aeroplane Co., Ltd. [i'Jllon, BrUtd

Proteus 706 PrOUII! 7M Olympll!!.BOt Orpbel.l8

D. Napier" Son Ltd. London W.3 Rolla-Royce Ltd.

D""

,..

...

52.SW

39.5" 395' '0'

2.650

Nomad 2

3."" 1,575

Elod

Derwent R. D. 7 NeJ:leR.N.2 Dart Mk.505 A"", R.A.2 Avoc.R.A.3 Avon R. A. 7 AVOD R. A. 7R'

1,23) 1,121

1.0\0

'.«IS

2,2402.-4~

2,9130*' 2,440 2,860 2,710 2,460

2.2

~ ~()

>

-!:

I\)

t~

WRIGHT

4~

WRIGHT

ENGINES - - - - - - - - - - -

4k

ENGINES - - - - - -_ _ _ __

t::l

~ ~z

ALTITUDE PERFORMANCE CURVE

no

MODEl 749C 18BO 1

COMP. RA

REO. GEAR RATIO ..4375:1

SUPER. RATIO 6.46 & 8.67:1

OPERATION AND INSTALLATION DATA

IMPELLER DIA. 13 in.

6.50: 1

fUEL GRADE 100/130

PERFORMANCE

At Standard Atmospheric Conditions with Operating Mixture Strengths Unless Otherwise Specified

(AlI1'atingJ are baJ(;!t/ on the

&l IIle

of lOV/}30 grade luel and operati.ng -mixture strengthI.) 951CI88DI UHP/RPM/ALl

Take'off (5 minmes maximum) Low Ratio High Ratio

2600

Rared Power (METO) !.ow Ratio High Ratio Cruise Power Low Ratio High Ra[io

!i?

F

149CI8BDI BHP/RPM/ AU

2500/280J/5L ro 3800

2500/2800/51. to 3800 1900/2600/11500 ro 15500

21OJ/2400/51.

.2100/2400/5L to 5500 1800/2400/9500 ro 16000

to

5500

1470/2300/SL to 13000

1470/2300/SL ro 13000 140012300/14500 ro 21000

2000

~ § ..... CD

~

::r' C/lc+

'""'"

!$

dl1;'

1:'1

","0

:>-

o '1

...,,,

WEIGHTS

","II>

~'"

2w '"o x:

ScanJarJ Engine Tllcal Dry Weight

1800

2869 Ib 2884 Jb

Standard Equipmem Included in Total Dry Weight

1600

w

1400

'" 1200

Optional Equipment and Resulting Variations from Standard Weight

r'TT',''';''',''';''';-'',''',''',''';''';''',''; .., '; ",', i ,': SL. 5 10

: ::

15

ALTITUDE-THOUSAND FEET

. : : ' , : ' : : :: . ,,,,I

"

20

25

Reduction gear, with .5625 ratio

~

::a

~ CIl

sa ~

0

~ ~

Approximate W.igh,

35 20 2

lb lb 1b

Approxima,e IlIcr.ase

Water injection, including control unit and discharge nozzles bur nor including regulator, tank, tUbing, etc.

30

0 0

tj

Manifold pressure rt'gu1awr Gear-driven fan, including afterbody and fan drive Propeller-speed fan, including afterbody Reduction gear, with .35 r.atio

200

Cl

" ..... ~

G~

Front Cylinder Exhaust Pipe Extensions Torquemerer Pressure Transmitter

400

t""

f;,J

!,J ;c,

Six Point Dynamic Suspension Assemblies

600

H

tv

Scintilla DLN-9 low-tension ignition system includ log magnero, discriburofS, and high-tension coils; also radio shielding and spark plugs

Standard Equipment Not Included in Total Dry Weight

til

C/l 'd t'l t'j

Supercharger drive (single-speed or two-speed as indica{(;d)

800

","r> 011>

-d

Bendix 58-18-B.)A fud iJljection system including master control, injector pumps and drives, pump supply and injection lines, .1Ild injecw[ nozzles

1000

0'"

"'c+ c+ ,..-

III

Provision for double-acting propeller piech control Torquemerer Priming sysrcm 011 engine Firf seal adapter tlange Cooling air deflectors Accessory drives Accessory drive covers Reducrion gear with A:':' 75 ratio

V)

:;'"

Model 951C18HD1 Model 749Cl8BOl

gj

g

15.8 168 80 10 3.8

1b Jb

lb Ib Jb

Approximate Oecrease

10

Jb

~ ~

- - - - - - - - - - - WRIGHT ESTIMATED SPECIFIC

t;-k

WRIGHT ENGINES - - - - - - - - - -

t;-k

ENGINES - - - - - - - - - - -

ALTITUDE PERFORMANCE CURVE-HIGH

FUEL CONSUMPTION CURVE-HIGH

RATIO

MODEl 749(18801

COMPo RATIO 6 ..50:1

IMPELLER: OIA. 13 in.

RED. GEAR RATIO .4375:1

SUPER. RATIO 8.67:1

fUel GRADE '00{130

RATIO

MODEL 749C18601

COMPo RATIO 6 ..50:1

IMPElLER OIA. 13 in.

RED. GEAR RATIO .4375:1

SUPER. RATIO 8.67:1

FUEL GRADE 100/130

At Standard Atmospheric Conditions with Operating Mixture Strenoths Unless Otherwise SpeciAed

~

;

At Standard Atmospheric Conditions with Operating Mixtur. Strenoths Unless Oth.rwi.e Specified

.88 2600

.86

:z:

~

.84 .82

t::t

:.-

2400

F

2200

H

..............

.80

~

g

.78 .76

'" J:

"-

ili

"'

1

:.ID

.66

IX

w

.. ... ~

3:

.68

u

iL

U w a..

V)

w

...""-


F. MINIMUM CLEARANCE ALLOW- ~M ) z ANCE FOR AIRCRAFT TIRES ;~ ) Clearance allowances between 'he tire and the adjacent parts of the aircraft should be based on the maximum overall tire dimensions shown In the tobles plus the growth due to service as calculated in "En, plus the increase in diameter due to centrifugal force. Minimum radial and lateral clearances are determined from the following chart:

u. S.

/"

RECOMMENDED INFLATION FOR LESS THAN MAXIMUM LOADS

CI

;.

Where .he aircraft tire is '0 be used at loads lell than the maximum shown in .his manual, It is suggested that reference be made to deflection curves for speciflc lizel to delermine the appropriate inflation. These curves are available on request. 0)

t"d

en

000

(II

'"'d

.::mc:+o~'dO

1D8~~~:1~ g,()~C+cttj

..... f}:;I 0"

'd ct

[D

• > .....

OJ ~ f-' 4ct~RJH)ct Z (1 PO • • O#(o.t1~

..

0

I»

~

I

t-3\.oJ

oP I

0> 0>

o:4l

,

i!~g8 Q

I)

81:.:1"0Ira e..u> >!~

P.

:t~f'

(I

~: ;t~ ;z b :1.1

I?A.~:=

.. "1

DI' 0 .... 01 "':>' .. ~rta~ P 0 U'

'g!! ~~ ~

.; t7't3 ~ 0 !l.iii·~ x Z

(1''''

Q) ...... t:S~

r:~ 3::

.....

: loads.

NOTE: The distribution of Figure 3-7 may

be used. § 3.216 Maneuvering loads. (al At maneuvering speed VI> assume a sudden defle.ction of the elevator control to the maximum upward deflection as limited by the control stops or pilot effort, whichever is critical. NOTE: The average loading of Figure 3-3 and the distribution of Figure 3-8 may be

used. In determining the resultant normal force coetl'lcient for the tall under these conditions, It will be jJ€rmlssible to assume that the angle of attack of the stabilizer with respect to the resultant direction of air !low Is equal to that Which occurs when the airplane is in steady unaccelera ted !light at a flight ~peed equal to VI>' The maximum elevator de!lection can then be determined from the above criteria and th~ tall normal No. 136--11

CONTROL SURFACE

n IS THE POSITIVE LIMIT MANEWERING LOAD FAc-TOR USED IN DESIGN.

NOTE: .. SHALt NoT 8E LESS THAN 12 PSF IN ANY CASE.

o

Horizontal tail surfaces. The

Balancing

AVERAGE MANEUVERING LOADINGS

LIMIT AVERAGE MANEUVERING LOADING: ii • K .. 4~ (PSF) WHERE

Correction noted at 14 F. R.

horizontal tail surfaces shall be designed for the conditions set forth In §§ 3.215-

Ailerons

w

HORIZONTAL TAIL SUl!FACES

3.214

Obtain was function of W /S and surface deflection in same manner as outllned In (1) above, use distribution of Figure 3-8; (4) Condition § 3.219 (b): Obtain 14 from Curve C, use distribution of Fi&llre 3-7; (5) Condlti6n § 3.219 (c): Obtain w from Curve A, use distribution of Figure 3-9. (Note that condition § 3.220 generally Will be more critical than this condition.)

and down directions. Use distribution of Figure 3-10.

§ 3.213 Tnm tab effects. The effects of trim tabs on the control surface design conditions need be taken into account only in cases where the surface loads are limited on the basis of maximum pilot effort. In such cases the tabs shalJ be considered to be deflected in the direction which would assist the pilot and the deflection shall correspond to the maximum expected degree of "out of trim" at the speed for the condition under consideration. §

Vertical Tail sur/aces (3) Condition § 3.219 (a):

(6) In lieu of conditions § 3.222 (b) : Obtain from Curve B, acting In both up

The Administrator will accept the following procedure as giving proper consideration of automatic pilot systems in assisting the pilot under § 3.212: The autopilot effort need not be added to human pilot effort but the autopilot effort shall be used for design if it alone can produce greater control surface loads than the human pilot. [12 F. R. 3436. 36J

4049

FEDERAL REGISTER

10

5W

15

(PSF)

FIG. 3-3(b)-UMIT AVERAGE MANEUVERING CONTROL SURFACE LOADING

2:

--_.J..-..,....-

~

I.O'!--.,...--t-+-----t.-.....

g~ .I!:2

.:~ (!)

zZ

,

~!:2 ::> 0 .&!-_ _ _ _+-____-/ ~t:1

IIJ

Z

':IE"

I-

~ ;:j I

.1

1. V

FIG. 3,4 -

.2

.3

OVERALL LENGTH OF AIRPLANE OESIGN SPEED (MPH)

.4

(ill

2

'-.&

MANEUVERING TAIL LOAD INCREMENT (UP OR DOWN)

......

~

li..

CJ)

0.

V!.!J.2S!.

l . _ . _ _ _ _ ' ' __

_~_

::J

CJ1 Q

-

C)

lao

z

is

« o

!.!2

..J

I~

~

CJ) (~

::>

.!!2

-p

100

~ o o w

w (!)

·t

a:: w ~

-------

«

(..!)

I. :

, J

;i Ii

10

15

W

~=(PSF)

Sv·

S

rn.

3.5to)-DOWN GUST LOADING ON HORIZONTAL TAIL SURFACE

'T

40

20

~O

c::: ~

NOTE: THESE CURVES ARE FOR JlSPECT RATIO R 05;·FO OTHEH ASPECT RATIOS MUL·{PLY LOADINGS

I

I

I~

o

H

ao

60

80

BY ..2lL1 (R+21 L I 100

~

I

I 0

.. -

~

!:d =)

MAXIMUM WEIGHT AREA OF VERTiCAL TAIL SURFACE

:Ill!

C

FIG. 3-6 -GUST LOADING ON VERTICAL TAIL SURFACE

r-

~

8

H 0

:l:fj5 ~

z

.

c ~

~iJ :!I. !::=== -E ·39-' I

~

~ ~------------r--r~-------r--------~--r-~------~~J 0.

C)

Z

o « o

..J

c

---'---l

_

.,

NOTES' (0) IN BALANCING CONDITIONS 03.2211 P ·40% OF NET BALANCING LOAD (FLAPS RETRACTED) P·O (FLAPS DEFLECTED. U,l IN CONDITION 03.2221 (b) po 20 % OF NET TAIL LOAD

FIG. 3-7 TAIL SURFACE LOAD DISTRIBUTION

1-

~ ooF-

=8 ~()I

(ACTING IN A

/DIRECTIDN OPPOSITE TO THE STABILIZER LOAD.)

~ .•

i w:'

,.

s;:

0

en

()I

5!ii z

~..~:~ P.'~RF;E LOAD DISTRIBUTION

(!)

f

0..

r----.~I L ___ ~~,--:-+-

;:)

W (!)

«

a::

w ~

__

5

,~

3-5(b) -

..

.....

OJ

W 10

5 FIG.

r;

III

15

20

a(PSF)

UP GUST LOADING ON HORIZONTAL TAIL SURFACE

FIG. 3-9 TAIL SURFACE LOAD DISTRIBUTION

FIG. 3-10 AILERON LOAD DISTRIBUTION

A6a: 16

CIVIL AIR REGUI.A.TIONS, PART 3 (GAR 3)

Saturday, July 16, 1949 force coefllclent can be obtained from the data given In NACA Report No. 688, "Aerodynamic Characteristics of Horizontal TaU Surfaces," or other appllcable NACA reports.

(b) Same as case (a) except that the elevator defiection is downward. NOTE: The average loading of Figure 3-3 and the distribution of Figure 3-8 may be

used.

(c) At all speeds above Vp the horizontal tail shall be designed for the maneuvering loads resulting from a sudden upward deflection of the elevator, foHowed by a downward defiection ot the elevator such that the following combinations of normal acceleration and angular acceleration are obtained: Condition

Airplane normal accelera· tion n

V-airplane speed In miles per hour, St= taU su.rface area In square feet, at ~ slope of 11ft curve of taU surface, CL per degree, corrected for aspect ratiO, G .. -slope of 11ft curve of wing, Ct per degree, S w=aspect ratIo of the wing.

§ 3.218 Unsymmetrical loads. The maximum horizontal tail surface loading (load per unit area), as determined by the preceding sections, shall be applied to the horizontal surfaces on one side of the plane of symmetry and the following percentage of that loading shall be applied on the opposite side: %=100--10 (n--1) where:

n Is the specified positive maneuverIng load factor.

AngulRr acreleration radian/sec)

In any case the above value shall not be greater than 80 percent. VERTICAL TAIL SURFACES

45

Down load ••••••••••

I. 0

+v n.. (n .. -l.5)

Up load. .......... ..

n.

-T' 71. (n .. -I.5)

45

where: nm=posltlve llmlt maneuvering load factor used In the design of the airplane. V=lnltial speed In mUes per hour.

(d) The total tail load for the conditions specifled in (c) shall be the sum of: (1) The balancing tail load corresponding with the condition at speed V and the specified value of the normal load factor n, plus (2) the maneuvering load increment due to the specified value of the angular acceleration. NOTE: The maneuvering loaa Increment of Figure 3-4 and the distributions of Figure 3--8 (for downloads) and Figure 3-9 (for uploadS) may be used. These distributions apply to the total tal! load. § 3.217 Gust loads. The horizontal tail surfaces shall be designed for loads occurring in the following conditions: (a) Positive and negative gusts of 30 feet per second nominal intensity at speed V c, corresponding to fiight condition § 3.187 (a) with flaps'retracted. NOTE: The average loadings of Figures 3-5 (a) and 3-5 (b) and the distribution of Figure 3-9 may be used for the total

tail loading In this condition.

(b) Positive and negative gusts of 15 teet per second nominal intensity at speed Vt, corresponding to flight condition § 3.190 (b) with flaps extended. In determining the total load on the horizontal tail for these conditions, the initial balancing tail loads shall first be determined for steady unaccelerated fiight at the pertinent design speeds Vc and Vt. The incremental tail load resulting from the gust shall then be added to the initial balancing tail load to obtain the total tail load. NOTE: The Incremental tall load due to the gust may be computed by the followlna formula: M=0.1 KUVStat

4051

FEDERAL REGISTER

(1- 3!:",)

where: ,:It=the limit gust load Increment on the tail in pounds; K=gust coefllclent K In § 3.188, U=nomlnal gust Intensity In feet per second,

Maneuvering loads. At all speeds up to VI>: (a) With the airplane in un accelerated fiight at zero yaw, a sudden displacement of the rudder control to the maximum defiectlo;l as limited by the control stops or pilot effort, whichever is critical, shall be assumed. § 3.219

NOTE: The average loading of Figure 3-$ and the distribution of Figure 3-8 may be used.

(b) The airplane shall be assumed to be yawed to a sideslip angle of 15 degrees while the rudder control is maintained at full deflection (except as limited by pilot effort) in the direction tending to increase the sideslip. NOTE: The average loading of Figure 3--3 and the distribution of Figure 3-7 may be used.

(c) The airplane shall be assumed to be yawed to a sideslip angle of 15 degrees while the rudder control is maintained in the neutral position (except as limited by pilot effort). The assumed sideslip angles may be reduced if it is shown that the value chosen for a particular speed cannot be exceeded in the cases of steady slips, uncoordinated rolls from a steep bank, and sudden failure of the critical engine with delayed corrective action. NOTE: Thp. average loading of Figure 3-$ and the distribution of Figure 3-9 may be used.

(a) The airplane shall be assumed to encounter a gust ot 30 feet per second nominal intensity, normal to the plane of symmetry while in unaccelerated fiight at speed V e• (b) The gust loading shall be computed by the following formula: § 3.220

Gust loads.

-

}(UVm

'" -

slope of 11ft curve of vertical surface, C L per radian, corrected for aspect

ratio, W= design weight In pounds, vertical surface area in square feet.

Sv -

(c) This loading applies only to that portion of the vertical surfaces having a well-defined leading edge. NOTE: The average loading of Figure 3-8 and the distribution of Figure 3--9 may be used. § 3.221 Outboard fins. When outboard fins are carried on the horizontal tail surface, the tail surfaces shall be designed for the maximum horizontal surface load in combination with the corresponding loads induced on the vertical surfaces by end plate effects. Such induced effects need not be combined with other vertical surface loads. When outboard fins extend above and below the horizontal surface, the critical vertical surface loading (load per unit area) as determined by §§ 3.219 and 3.220 shall be applied: (a) To the portion of the vertical surfaces above the horizontal surface, and 80 percent of that loading applied to the portion below the horizontal surface, . (b) To the portion of the vertical surtaces below the horizontal surface, and 80 percent of that loading applied to the portion above the horizontal surface. AILERONS, WING FLAPS, TABS, ETC.

Ailerons. (a) In the symmetrical fiight conditions (see §§ 3.1833.189), the ailerons shall be designed for all loads to which they are subjected while in the neutral position. (b) In unsymmetrical flight conditions (see § 3.191 (a) ), the ailerons shall be designed for the loads resulting from the following deflections except as limited by pilot effort: (1) At speed V p it shall be assumed that there occurs a sudden maximum displacement of the aileron control. (Suitable allowance may be made for control system deflections.) (2) When Vc is greater than V p , the aileron deflection at Ve shall be that required to produce a rate of roll not less than that obtained in condition (1). (3) At speed Va the aileron deflection shall be that required to produce a rate of roll not less than one-third of that which would be obtained at the speed and aireron deflection specifled in condition § 3.222

(1).

NOTE: For conventional allerons. the deflections for conditions (2) and (3) may be computed from: _ Vp

Il,-

.

ve"h,

0.5 Vp

an d 8,=--8 ,: Va

where: cS,=total alleron defiectlon (sum of both aileron defiections) in condition (1) •

W=---

K= 1.33 (W /S.) .,,' except that K shall

a,=total allercn defiection In condition (2). Il.=total defiectlon In condition (3). In the equation for 0.1. the 0.5 factor Is used instead of 0.33 to allow for wing torsional fiexiblllty.

not be less than 1.0. A value of K obtained by rational determination may be used. U= nominal gust intenSity In feet per second, V= airplane speed In miles per hour,

(c) The criticalloadin€, on the ailerons should occur in condition (2) if Va is less than 2Vc and the wing meets the torsional stiffness criteria. The normal force coeIDcient eN for the ailerons may be taken as 0.048, where 8 is the defiec-

575

where: W=

average limIt unIt pressure pounds per square foot,

In

4.5

CIVIL AIR REGULATIONS, PART 3 (CAR 3) 4052

Aoa: 17

RULI!S AND REGULATIONS

tion of the indiVidual aileron in degrees. The critical condition for wing torsional loads will depend upon the basic airfoil moment coefficient as well as the speed, and may be determined as follows: T. T.

(C m -.01831)VI (Cm-.0182t1 Ve"

where:

T.IT. is the ratio of wing torsion In

condition (b) (3) to that in condition (b)

(2).

and 831 are the down defiect~()n~ 01 the Individual aUeron in condlt1ons (b) (2) and (3) respectively.

1I21

(d) When T,IT, is greater than 1.0 condition (b) (3) is critical; when Ta/T. is less than 1.0 condtion (b) (2) is critical. (e) In lieu of the above rational conditions the average loading of Figure 3-3 and the distribution of Figure 3-10 may be used. § 3.223 Wing !taps. Wing fiaps, their operating mechanism, and supporting structure shull be designed for critical loads occuning in the fiap-extended flight conditions (see § 3.190) with the fiaps extended to any position from fully retractea to fully extended; except that when an automatic fiap load Err..it:~g d2vice is employed these parts may be designed for critical combinations of air speed and fiap position permitted by the device. (Also see §§ 3.338 and 3.339,) The effects of propeller slipstream corresponding to take-off power shall be taken into account at an airplane speed of not less than 1.4 V, where V, is the computed stalling speed with fiaps fully retracted at the design weight. For investigation of the slipstream condition, the airplane load factor may be assumed to be 1.0. § 3.224 Tabs. Control surface tabs shall be designed for the most severe combination of air speed and tab defiection likely to be obtained within the limit V-n diagram (Fig. 3-1) for any usable loading condition of the airplane. § 3.225 Special devices. The loading for special devices employing aerodynamic surfaces, such as slots and spOilers, shall be based on test data. CON"rROL SYSTEM LOADS

§ 3.231 Primary !tight controls and 811stems. {a) Flight control systems and

supporting structures shall be designed for loads corresponding to 125 percent. of the computed hinge moments of the movable control surface in the conditions prescribed in §§ 3.211 to 3.225, subject to the following maxima and minima: (1) The system limit loads need not exceed those which can be produced by the pilot and automatic devices operating the controls. (2) The loads shall in any case be sufficient to provide a rugged system for service use, including consideration of jamming, ground gusts, taxying tail to wind, control inertia, and friction. (b) Acceptable maximum and minimum pilot loads for elevator, aileron, and rudder controls are shown in Figure 3-11. These pilot loads shall be assumed to act at the appropriate control grips or pads in a manner simulating flight con-

ditions and to be reacted at the attachments of the control system to the control surface horn. § 3.231-1

Hinge moments (CAA pol-

fcies which apply to § 3.231). The 125 percent factor on computed hinge moments provided in § 3.231 (a) need be applied only tb elevator, aileron and rudder systems. The Administrator will accept a factor as low as 1.0 when hinge moments are based on test data, the exact reduction which the Administrator will accept, depending to an extent upon the accuracy and reliability of the data. [12 F. R. 8436. 36]

Correction noted at 14 F. R.

System limit loads (cAA policies which apply to § 3.231 (a) (1». § 3.231-2

The Administrator will accept the followinb procedure as compliance with

!!ir:: i~)co~~n~~~n wti~~ ~~~oh~~a~ pilot, the autopilot effort need not be added to human pilot effort but the autopilot effort shall be used for design if it alone can produce greater control,surface loads than the human pilot. When the hUman pilot acts in opposition to the autopilot, that portion of the system between them shall be designed for the maximum effort of human pilot or autopilot, whichever is the lesser. [12 F. R. 3436. 86]

Correction noted at 14 F. R.

§ 3.232 Dual controls. When dual controls are provided, the systems shall be designed for the pilots operating in opposition, using individual pilot loads equal to 75 percent of those obtained In accordance with § 3.231. except that the indiVidual pilot loads shall not be less than the minimum loads specified in Fig.ure 3-11. § 3.233

Ground gust conditions.

(a)

The follOWing ground gust condition:; shall be investigated in cases where a deviation from the specific values for minimum control forces listed in Figure 3-11 is applicable. The following conditions are intended to simulate the lo~.d­ ings on control surfaces due to ground gusts and when taxying with the wind. (b) The limit hinge moment H shall be obtained from the following formula:

where I H-limlt hinge mOlI1ent (foot-pounds). c-mean chord of the control surfaa. aft of the hinge Hne (teet). 8=area of control surface aft of the hlnie line (square feet). q=dynamic pressure (pounds per square foot) to be based on a deSign speed not less than 10VWIS+10 miles per hour, except that tlul design speed need not exceed 00 mlles per hour. K=factor 8.8 Ipec1.fied belowl

Sur/ace (a) Aileron ____________________ Control column locked or lashed in mid-pOSition. (b) Alleron ____________________ Ailerons at full throw; + moment on one alleron, - moment on the other. (c) (d) Eaevator _______________ Elevator (c) full up (-). and (d) full down (+). (e) (f) Rudder ________________ Rudder (e) in neutral, and (f) at tull thtow.

K +0.70 ±O.50

li:0.75 ±0.70

(c) As used in paragraph (b) in connection with ailerons and elevators, a positive value of K indicates a moment tending to depress the surface while a negative value of K indicates a moment tending to raise the surface. § 3.233-1 Ground gust loads (cAA policies which apply to § 3.233). Section 3.233 requires ground gust loads to be in-

vestigated when a reduction in minimum pilot effort loads is desired. In such cases the entire system shall be investigated for iround gust loads. However, in instances where the designer desires to investigate ground gust loads without intending to reduce pilot effort loads, the ground gust load need be carried only from the control surface horn to the nearest stops or gust locks, including the stops or locks and their supporting structures. [12 F. R. 8436. 36]

Correction noted at 14 F. R.

§ 3.234 Secondary controls and systems. Secondary controls, such as wheel

brakes, spoilers, and tab controls, shall be designed for the loads based on the maximum which a pilot is likely to apply to the con trol in question.

LIMIT PILOT LOAD!

Control

Maximum loads for design weight Wequal to or less than ~,OOO lbs.'

Alleron: Stick _____________________ •• __ • _____ _ 67 pounds _______________________________ • _____ ._ Wheel , ______ • __ •• _._._. _____ • ____ • __ 53 Din-pounds , _________ • ____________________ ._ Elevator: Stick _______________________________ _ 167 pounds ___________________________________ • __ 200 pounds _______________ • ______________________ 200 pounds ______ • _________ ._. _________ ._. _______

Rud'X;~~~~~:~~:~:~::::~::::::~~~~:::::~

Minimum load! I

40 pounds. 40 Din-pound •• 100 pounds. 100 pounds. 130 pounds.

I For design weight W greater than 5.000 pounds the above specified maximum values sball be increased linearly witb weigbt ~o 1.5 tim~s the specified values at "design weigbt of 25,000 pounds. • If t?e deSIgn of any individual set of control systems or surfaces is sucb as to make these specified minimum loads inapplicable, values corresponding to tbe pertinent binge moments obtained according to p.233 may be used instead, except tbat m any case values less than 0.6 of tbe specified minimum loads shall not be employed. 'Tbe critICal portIOns of tbe aileron control system sball also be designed for a single tangential force baving a limit Talue rqual to 1.25 times tbe couple force determined from tbe above criteria. 'D-wbeal diameter.

FIG. 3-11-PILOT CONTROL FORCE LIMITS

A6a:18 Saturday, July 16, 1949

cmL AIR REGULATIONS, PJlRT 3 (CAR 3) FEDERAL UGfSTER

GROUND LoADS

§ 3.241 Ground loads. The loads specified in the following conditions shall be considered as the external loads and inertia forces which would occur in an airplane structure if it were acting as a rigid body. In each of the ground load conditions specified the external reactions shall be placed in equilibrium with the linear and angular inertia forces in a rational or conservative manner. § 3.242 Design weight. The design weight used in the landing conditions shall not be less than the maximum weight for which certification is desired: Provided, however, That for multiengine airplanes meeting the one-engine-inoperative climb requirement of § 3.85 (b), the airplane may be designed for a design landing weight whil:h is less than the maximum design weight, if compliance is shown with the following sections of Part 4b in lieu of the corresponding reqUirements of this part: the ground load requirements of § 4b.241, and shock absorption requirements of § 4b.371 and its related sections, the wheel and tire requirements of §§ 4b.391 and 4b.392, and the fuel jettisoning system requirements of § 4b.536. § 3.243

Load factor for landing con-

In the following landing conditions the limit vertical inertia·load factor at the center of gravity of the airplane shall be chosen by the designer but shall not be less than the value which would be obtained when landing the airplane with a descent velocity, in feet per second, equal to the following value: ditions.

V =4.4 (W IS) )i

except that the descent velocity need not exceed 10 feet per second and shall not be less than 7 feet per second. Wing lift not exceeding two-thirds of the weight of the airplane m88 be assumed to exist throughout the landing impact and may be assumed to act through the airplane center of gravity. When such wing lift is assumed, the ground reaction load factor may be taken equal to the inertia load factor minus the ratio of the assumed wing lift to the airplane weight. (See § 3.354 for reqUirements concerning the energy absorption tests which determine the limit load factor corresponding to the required limit descent veloCities.) In no case, however, shall the inertia load factor used for design purposes be less than 2.67, nor shan the limit ground reaction load factor be less than 2.0, unless it is demonstrated that lower values of limit load factor will not be exceeded in taxying the airplane over terrain having the Il}aximum degree of roughness to be expected under intended service use at all speeds up to take-off speed.

tions which are considered to conform with §§ 3.245-3.247.) § 3.245 Levellanding-(a> Tail wheel type. Normal level fiight attitude. (b) Nose wheel type. Two cases shall be considered: (1) Nose and main wheels contacting the ground simultaneously, (2) Main wheels contacting the ground, nose wheel just clear of the ground. (The angular attitude may be assumed the same as in subparagraph (1) of this paragraph for purposes of analysis.) (c) Drag components. In this condition, drag components simulating the forces required to accelerate the tires and wheels up to the landing speed shall be properly combined with the corresponding instantaneous vertical ground reactions. The wheel spin-up drag loads

may be based on vertical ground reactions, assuming wing lift and a tiresliding coemcient of friction of 0.8, but in any case the drag loads shall not be less than 25 percent of the maximum vertical ground reactions neglecting wing lift. § 3.245-1 Wheel spin-up loads (CAA policies which apply to § 3.245).

Tail wheel type Condition

Level landing

Reference section •• _••••••••••••••••••••••••• Vertical component at c. g.•...••..•••.•••••• Fore and aft component at c. g ............... Lateral oomtnent in either direction at c. g. Shock absor r extension (hydraulic shock absorber) .....•.......••......••..•.......• Shock absorber deflection (rubber or spring shock absorber) ••....•••••• " •• ' •• '" ...... Tire deflection••••••.••.•..••...•••••••••.... Main wheel loads (both WheeIS) ••• _ •••.

~

Nose wheel type.

Tail-down landing

3.245 (8) I 3.246 (8)

Levellandlng Le vel landing nose wheel TaIl-dOWll with inclined with just clear 01 landinc reactions ground ~

3.245 (b) (l)

13.245 (b) (2) 13.246(b)(c)

nW KnW

nW 0 0

nW KIIW 0

nW KnW 0

nW

Note (2)

Note (2)

Note (2)

Note (2)

'Note (2)

100%

100%

100%

100% Static nW KV,

Statio

a

Static nW KV, 0 Tail (nose) wheel loads •.••••••••••..••.• 0 Notes ••••••••••••••••••••••••••••••••••••••• (1) and (3)

{b': {b';

Static nWb/d 0 IIW8/d 0

Static nWb'/d' KV, IIWa'/d' Kl;,

-.....---_ .... _..

0 0

0 0 (1) and (3)

(1)

100%

.w 0 0 0

(3)

NOTE (1).-K may be determined as follows: K-O.25 (or W-3,OO9 pounds or less; K-O.33 for W-6.000 pounds or greater, with linpar variation of K between these weights. NOTE (2).-For tbe purpoS section 5.00 O-f ANC-5,

Amendment No. I, allowable design property c.olumns headed "Army-Navy" represent design properties which will be equalled or exceeded by the properties possessed by approximately 00 percent of the material. All other allowable design property columns relate to the minimum guaranteed properties and are based on values given in the various material specifications. The Administrator will permit uses of these design properties as outlined in subparagraphs (1) and (2) of this section, based on the objectives of § 3.301. (1) In the case of structures where the applied loads are eventually distributed through single members within an assembly, the failure of which would result in the loss of the structural integrity of the component involved, the guaranteed minimum design mechanical properties .u~.c;u iH .nl'lC-5 shaii ue u~t:d.

NOTE: Typical examJllea of such items are: 1. Wing lift struts. 2. Spars in two-spar w1nga.

3. Sparcaps In regions such III! wing cutouts and wing center sections where loads are transmitted through caps only. 4. Primary attachment fittings dependent on single bolts for load transfer. (2) Redundant structures wherein partial failure of individual elements would result in the applied load being safely distributed to other load carrying members, may be designed on the basis of the "90 percent probability" allowable.

NOTE: Typical examples of such Items are: 1. Sheet-stiffener combinations. 2. Multi-rivet or multiple bolt connections.

(b) Certain manufacturers have indicated a desire to use design value greater than the guaranteed minimums even in applications where only guaranteed minimum values would be permitted under paragraph (a) of this section, and have advocated that such allowables be based on "premium selection" of the material. Such increased design allowabIes will be acceptable to the Administrator: Provided, That a specimen or specimens of each individual item are tested prior to its use, to determine that the actual strength properties of that particular item will equal or exceed the properties used in design. This, in effect, results in the airplane o.~ materials manufacturer guaranteeing higher minimum properties than those given in the basic procurement specifications. (c) When strength testing is employed to establish design allowables (such as in the c,ase of sheet-stiffener compression tests), the test results shall be reduced to values which would be met by material having the design allowable material properties for the part under consideration, as covered in subparagraphs (1) and (2) of this sectwn. Non:: Sections 1.543 and 1.544 of ANC-5 outline two means of accomplishing this, but are by no means considered as the only methods available. [12 F. R. 3436. Correction noted at 14 F. R. 36] § 3.302 SpeCial factors. Where there may be uncertainty concerning the actual strength of particular parts of the structure or where the strength Is likely to deteriorate in service prior to normal replacement, increased factors of safety shall be provided to insure that the reliability of such parts is not less than the rest of the structure as specified in § § 3.303-3.306. § 3.303 Variability factor. For parts whose strength is subject to appreciable variability due to uncertainties in manufacturing processes and inspection methods, the factor of safety shall be increased sufficiently to make the probability of any part being under-strength from this cause extremely remote. Minimum variability factors (only the highest pertinent variability factor need be considered) are set forth in § § 3.3043.306. § 3.304 Castings. (a) Where visual inspection only is to be employed, the variability factor shall be 2.0. (b) The variability factor may be reduced to 1.25 for ultimate loads and

A5a:22 Saturday. July 16. 1949 1.15 for limit loads when at least three sample castings are tested to show compliance with these factors, and all sample and production castings are visually and radiographically inspected in accordance with an approved inspection specification. (c) Other inspection procedures and variability factors may be used if found satisfactory by the Administrator. § 3.304-1 Casting factors (CAA policies which apply to § 3.304). With reference to paragraphs (b) and (c} of § 3.304, the Administrator has approved specific proposals which permit the U3e of lower casting factors as speCified in (b), with 100 percent radiographic inspection on initial runs, but with radiographic inspection gradually reduced on production lots as it becomes evident that adequate quality control has been established. All such procedures require the submittal and execution of a 'Satisfactory process speCification and statistical proof tnat adequate quality control has been achieved. [12 F. R. 3437. Correction noted at 14 F. R. 36] § 3.305 Bearing factors. (a) The factor of safety in bearing at bolted or pinned joints shall be suitably increased to provide for the following conditions: (1) Relative motion in operation (control surface and system joints are covered in §§ 3.327-3.347). (2) Joints with clearance (free fit) subject to pounding or Vibration. (b) Bearing factors need not be applied when covered by other special factors. § 3.306 Fitting factor. Fittings are defined as parts such as end terminals used to join one structural member to another. A multiplying factor of safety of at least 1.15 shall be used in the analysis of all fittings the strength of which is not proved by limit and ultimate load tests in which the actual stress conditions are simulated in the fitting and the surrounding structure. This factor applies to all portions of the fitting, the means of attachment. and bearing on the members joined. In the case of integral fittings, the 'part shall be treated as a fitting up to the point where the section properties become typical of the member. The fitting factor need not be applied where a type of joint design based on comprehensive test data is used. The following are examples: continuous joints in metal plating, welded jOints, and scarf joints In wood, all made in accordance with approved practices. § 3.307 Fatigue strength. The structure shall be designed, insofar as practicable, to avoid points of stress concentration where variable sttesses above the fatigue limit are likely to occur in normal service. FLUTTER AND VIBRA'rION

3.311 Flutter and vibration prevention measures. Wings, tail. and control surfaces shall be free from flutter. airfoil divergence, and control reversal from lack of rigidity, for all conditions of operation within the limit V-n envelope. and the following detail requirements shall apply: §

~o.136--1~

CIVIL AIR R,i!;~LATIONS. PAn;:

oJ

'-:'A..:':' J)

4051

FEDERAL REGISTER (a) Adequate wing torsional rigidity shall be demonstrated by tests or other methods found suitable by the Administrator. (b) The mass balance of surfaces shall be such as to preclude fiutter. (c) The natural frequencies of all main structural components shall be determined by vibration tests or other methods found satisfactory by the Administrator. WINGS

3.317 Proof of strength. The strength of stressed-skin wings shall be substantiated by load tests or by combined structural analysis and tests. § 3.318 Ribs. (a) The strength of ribs in other than stressed-skin wings shall be proved by test to at least 125 percent of the ultimate loads for the most severe loading conditions, unless a rational load analysis and test procedure is employed and the tests cover the variability of the particular type of construction. (b) The effects of ailerons and high lift devices shall be properly accounted for. Rib tests shall simulate conditions in the airplane with respect to torsional rigidity of spars, fixit3 conditions, lateral suPPOrt, and attachment to spars. § 3.318-1 Rib tests (CAA policies which apply to § 3.318). Section 3.318 was drafted so as to allow the proof of strengt:l of ribs in stressed skin wings to be made as part of lOO-percent ultimate load test of the wings, in cases where the complete wing is tested in such a manner as to simulate the actual air load distribUtion. In such cases the Administrator will not require that separate rib tests be made. When ribs of stressed skin wings are tested separately from the wing and a rational load distribution Is made, a suitable variability factor (see § 3.303) shall be employed in determining the test loads. Although no specific value is stated in § 3.303, a factor of 1.15 is considered acceptable. However, consideration may be given to a lower factor if SUch lower factor were substantiated by tests on a large number of ribs. §

[12 F. R. 3437. Correction noted at 14 P. R. 36] § 3.319 External bracing. When wires are used for external lift bracing they shall. be double unless the design provides for a lift-wire-cut condition. Rigging loads shall be taken into account in a rational or conservative manner. The end connections of brace wires shall be such as to minimize restraint against bending or vibration. When brace struts of large fineness ratio are used, the aerodynamic forces on such struts shall be taken into account. § 3.320 ·Covering. Strength tests of fabric covering shall be required unless approved grades of cloth. methods of support, attachment. and finishing are employed. Special tests shall be required when it appears necessary to account for the effects Of unusually high design air speeds, slipstream velocities. or other unusual conditions. § 3.320-1 Aircraft fabric (CAA rules which apply to § 3.320). See §§ 4b.302-1 and 4b.302-2 of this chapter.

[13 P. R. 7723]

CONTROL SURFACES (FIXED AN:: MOVABLE)

§ 3.327 Proof of strength. Limit load tests of control surfaces are required. Such tests shall include the horn or fi1lting to which the control system is attached. In structural analyses, tlgging loads due to wire bracing shall be taken into account in a rational or conservative manner. § 3.328 Installation. Movable taU surfaces shall be so installed that there is no interference between the surfaces or their bracing when each is held in its extreme positioq and all others are operated through their full angular movement. When an adjustable stabilizer is used, stops shall be provided which, in the event of failure of the adjusting mechanism, will limit its travel to a range permitt¢g safe fiight and landing. § 3.329 Hinges. Control sur f ace hinges, excepting ball and I()ller bearings, shall incorporate a mu~tiplying factor of safety of not less than 6.67 with respect to the ultimate bearing strength of the softest material used as a bearing. For hinges incorporating ball or roller bearings, the approved rating of the bearing shall not be exceeded. Hinges shall provide sufficient strength and rigidity for loads parallel to the hinge line. CONTROL SYSTEMS

3.335 General. All controls shall operate with sumcient ease, smoothness, and positiveness to permit the proper performance of their function and shall be so arranged and identified as to provide convenience in operation and prevent the possibility of confusion and subsequent inadvertent operation. (See § 3.384 for cockpit controls') § 3.336 Primary /tight controls. (a) nrl~"'.y :flIght controls are defined as those used by the pilot for the immediate control of the pitching, rolling, and yawing of the airplane. (b) For two-control airplanes the design shall be such as to minimize the likelihood of complet~ loss of the lateral directional control in the event of failure of any connecting or transmitting element in the control system. § 3.337 Trimming controls. Proper precautions shall be taken against the possibility of inadvertent, improper. or abrupt tab operations. Means shall be provided to indicate to the pilot the direction of control movement relative to airplane motion and the position of the trim device with respect to the range of adjustment. The means used to indicate the direction of the control movement shall be adjacent to the control, and the means used to indicate the position 01 the trim device shall be easily visible to the pilot and so located and operated as to preclude the possibility of confusion. Trimming devices shall be capable of continued normal operation notwithstanding the failure of anyone connecting or transmitting element in the primary flight control system. Tab controls shall be irreversible unless the tab is properly balanced and posseMes no unsafe fiutter characteristics. Irreversible tab systems shall provide adequate §

CIVIL AIR REGULATIONS, PART 3 (CAR 3) 4058 rigidity and reliability in the portion of the system from the tab to the attachment of the irreversible unit to the airplane structure. § 3.338 Wing f/.ap controls. The controls shall be such that when the flap bas been placed in any position uPon which compliance With the performance requirements is based, the flap will not move from that position exeept upon fUrther adjustment of the control or the automatic operation of a flap load limitIng device. Means shall be provided to !ndi~te the flap position to the pilot. If any flap position other than fully retracted or extended is used to show compliance with the performance reQUirements, such means shall indicate each such position. The rate of movement of the flaps in response to the operation of the pilot's control, or of an automatic device shall not be SUCh as to. result in unsatisfactory flight or performance characteristics under steady or changing conditions of air speed, engine power, and airplane attitude. (See § 3.109 (b) and (c).) § 3.339 Flap interconnection. (a,) The

motion of fiaps on opposite sides of the plane of symmetry shall be synchroniZed by a mechanical interconnection, unless the airplane is demonstrated to have safe flight characteristics while the flaps are retracted on one side and extended on the other. (b) Where an interconnection is used, In the case of multiengine airplanes, it shall be designed to account for the unsymmetrical loads resulting from flight with the engines on one side of the plane of symmetry inoperative and the remaining engines at take-off power. For single-engine airplanes, it may be assumed that 100 percent of the critical air load acts on one side and 70 percent on the other. § 3.340 stops. All control systems .shall be provided with stops which positively limit the range of motion of the control surfaces. Stops shall be so located in the system that wear, slackness, or take-up adjustments will not appreCiablY affect the range of surface travel. Stops shall be capable of withstanding the loads corresponding to th~ design conditions for the control system. § 3.341 Control system locks. When a device is provided for locking a control surface while the airplane is on the ground or water: (a) The locking device shall be so installed as to provide unmistakable warnIng to the pilot when it is engaged, and (b) Means shall be provided to preclude the possibility of the lock becomil;l.g engaged during Uight. § 3.342 Proo! of strength. Tests shall be conducted to prove compliance with limit load requirements. The direction of test loads shall be such as to produce. the most severe loading of the control system structure. The tests shall include all flttings, pulleys, anI! brackets used to attach ihe control system to the primary structure. Analyses or individual load tests shall be conducted to demonstrate compliance with the multiplying factor of safety reqUire-

A6a:23

RULES AND REGULATIONS

ments specified for control system joints subjected to angular motion. § 3.343 Operation test. An operation test shall be conducted by operating the controls from the pilot compartment With the entire system so loaded as to correspond to the limit air loads on the surface. In this test there shall be no jamming, excessive friction, or excessive deflection. CONTROL SYSTEM: DETAILS

General. All control systems and operating devices shall be so deSigned and installed as to prevent jamming, chafing, or interference as a result of inadequate clearances or from cargo, passengers, or loose objects. Special precautions shall be provided in the cockpit to prevent the entry of foreign objects into places where they might jam the controls. ProVisions shall be made to prevent the slapping of cables or tubes against parts of the airplane. § 3.345 Cable systems. Cables, cable fittings, turnbuckles, splices, and pulleys shall be in accordance with approved specii1cations. Cables smaller than 'nIinch diameter shall not be used in primary control systems. The design of cable systems shall be such that there will not be hazardous change in cable tension throughout the range of travel under operating conditions .and temperature variations. Pulley types and sizes shaY correspond to the cables with which they are used, as specified on the pulley specification. All pulleys shall be provided with satisfactory guards which shall be closely fitted to prevent the cables becoming misplaced or fouling, even when slack. The pulleys shall lie in the plane passing through the cable within such limits that the cable does not rub against the pulley flange. Fairleads shall be so installed that they are not required to cause a change in cable direction of more than 3 degrees. Clevis pins (excluding those not subject to load or motion) retained only by cotter pins shall not be employed in the control system. Turnbuckles shall be attached to parts having angular motion in such a manner as to prevent positively binding throughout the range of travel. Provisions for Visual inspection shall be made at all fairJeads, pulleys, terminals, and turnbuckles. § 3.344

§ 3.345-1 Cables in primary control systems (CAA interpretations which apply to § 3.345). Section 3.345 provides

that "cables smaller than %-inch diameter shall not be used in primary control systems." Primary control systems are normally considered to be the aileron, rudder, and elevator control systems. Hence this minimum of % inch need not be applied to tab control cables having high strength margins. However, in cases where the airplane would not be safely controllable in flight and landing with tabs in the most adverse positions required for the various critical trim, weight, and center of gravity conditions, the Administrator will require that tab syStems be so designed as to provide reliability equivalent to thRt required for primary systems. Examples are pulley

sizes, guards, use of fairleads, inspection provisiOns, etc. [1:1 P. R. 3437. Correction noted at 14 F. R. 86]

§ 3.346 Joints. Control system joints subject to angular motion in push-pull systems, excepting ball and roller bearing systems, shal! incorporate a multiplying factor of safety of not less than 3.33 with respect to the ultimate bearing strength of the softest material used as a bearing. This factor may be reduced to 2.0 for such joints in cable control systems. For ball or roller bearmgs the approved rating of the bearing shall not be exceeded. § 3.347 Spring devices. The reliability of any spring devices used in the control system shall be established by tests simulating serviee conditions, unless it is demonstrated that failure of the spring will not cause flutter or unsafe flight characteristics. LANDING GEAR SHOCK ABSORBERS

Shock absorbing elements in main, nose, and tail wheel units shall be substantiated by the tests specifled in the following section. In addition, the shock absorbing ability of the landing gear in taxying must be demonstrated in the operational tests of § 3.146. § 3.351

Tests.

§ 3.352

Shock absorption tests.

(a)

It shall be demonstrated by energy absorption tests that the limit load factors selected for design in accordance with § 3.243 will not be exceeded in landings with the limit descent velocity specified in that section. (b) In addition, a reserve of energy absorption shall be demonstrated by a test in which the descent velocity is at least 1.2 times the limit descent velocity. In this test there shal! be no failure of the shock absorbing unit, although yielding of the unit will be permitted. Wing lift equal to the weight of the airplane may be assumed for purposes of this test. § 3.352-1 Landing gear drop tests (CAA poliCies which apply to § 3.352).

(a) The following method has been approved by the Administrator for determining the effective mass to be dropped in drop tests of nose wheel landing gear assemblies pursuant to § 3.352 (a): For aircraft with nose wheel type gear, the effective mass to be used in free drop test of the nose wheel shall be determined from the formulp, for We (§§ 3.353 and 3.355) USing W=Wn where Wn is equal to the vertical components of the resultant force acting on the nose wheel, computed under the following assumptions: (1) the mass of the airplane concentrated at the center of gravity and exerting a force of 1.0 g downward and 0.33 g forward, (2) the nose and the main gears and tires in static position, and (3) the resultant reactions at the main and nose gears acting through the axles and parallel to the resultant force at the airplane center of gravity. NOTE: By way of explanation, the use ot an inclined reactions condition as the basis for determining the mass to be dropped with a nose wheel unit 11; based on rational dynamic investigation of the landing condition, as-

AOa:24

CIVIL AIR REUUL.riTIONS, PART 3 (CAR 3)

Saturday, July 16, 1949 sumlng the landing Is made with slmultaueous three-point contact, zero pitching velocity, and a drag component representing the average wheel spin-up reactions during the landing impact. Although spin-up loada on small alrplanes may be less than the value implied by the formUla, such aiI'plane. are more llkely to be landed with a nosing down pitching velocity, or In soft ground. The vertical component of the ground reaction Is specified above because the method of defining the direction of the inertia force at the center ot gravity gives a resultant effective mass greater than that of the airplane.

(b) The following procedure has been approved by the Administrator for determining the attitude in which the landing gear unit should be dropped pursuant to § 3.352 (a): The attitude in which a landing gear unit is dropped shall be that which simulates the airplane landing condition which is critical from the standpoint of energy to be absorbed by the particular unit, thus: (1). For nose wheel type landing gear, the nose wheel gear shall be drop tested In an attitude which simulates the three point landing inclined reaction condition; (2) the attitude selected for main gear drop tests shall be that which simulates the two-wheel level landing with inclined reactions condition. NOTE: In addition, It Is recommended that the main gear be dropped In an attitude simulating the taU-down landing with vertical reactions condition If the geometry of the gear Is such that this condition Is like1y to result In shock strut action appreciably different from that obtained In level attitude drop tests; for example, when a cantilever shock strut has a large Inclination with respect to the direction of the ground reaction. (3) Tail wheel units shall be tested in such a manner as to simulate the taildown landing condition (three-point contact). Drag components may be covered separately by the tail wheel "obstruction" condition. (c) The Administrator has accepted the following procedure for determining slopes of inclined platforms when such are used in drop tests: When the arbitrary drag components given on Fig. 3-12 (a) of this part are used for the design of the landing gear in the level landing conditions, the drag loads in the drop tests for these conditions may be simulated by dropping the units onto Inclined platforms so arranged as to obtain the proper direction of the resultant ground reactions in relation to the landing gear. (If wheel spin-up loads for these conditions are determined by rational methods and found to be more severe than the arbitrary drag loads, it is suggested that the spin-up loads be simulated by dropping the gear onto a level platform with wheels spinning,) In at least o11e limit drop test the platform should simulate the friction characteristics of paved runways and the rotational speed of the wheel just prior to contact should correspond to an airplane ground speed of 1.2 V,O' It is suggested that additional limit drops be made onto surfaces of lower friction coefficient and at several wheel rotational speeds; coefficients for example, corresponding to 0.6, 0.8 and 1.0 V' o'> The direction of wheel rotation in the drop test should be opposite to that which would

4059

FEDERAL REGISTER occur in landing the airplane. Spin-up loads which are slightly greater than the arbitrary drag loads can probably be simulated satisfactorily by 1nclined platforms, but platforms having greater InclinatIons may not simulate spin-up loads correctly and are not recommended. [12 P. R. 3437. Correction noted at 14 F. R. 36)

§ 3.353 Limit drop tests. (a) If compliance with the specified limit landing conditions of § 3.352 (a) is demonstrated by free drop tests, these shall be conducted on the complete airplane, or on units consisting of wheel, tire, and shoc~ absorber in their proper relation, from free drop heights not less than the following: h (Inches) -3.6 (WIS)G.I

except that the free drop height shall not be less than 9.2 tnches and need not be greater than 18.7 inches. (b) In simulating the permissible wing lift in free drop tests, the landing gear unit shall be dropped with an effective mass equal to:

d]

W =W[h+(1-L} , h+d where W. = the effective weight to be used in the drOp test. h=speclfied height of drop In Inohes. d=deflection under Impact of the tire (at the approved inflation pressure) plus the vertical component of the axle travel relative to the drop mass. The value of d used In the computation of W. shall not exceed the value actually obtained in the drop tests. W=W M tor main gear units, and shall be equal to the static weight on the particular unit with the airplane In the level attitude (with the nose wheel clear, in the case Of nose wheel type airplanes). W = W T tor tall gear units, and shan be equal to the static weight on Lhe taU unit with the airplane In the tall down attitude. W= W N for nose wheel units, and shall be equal to the static reaction which will exist at the nose wheel when the mass ot the airplane Is concentrated at the center of gravity and exerts a force ot 1.0g downward and a.33g forward. L=rat!o ot assumed wing 11ft to airplane weight, not greater than 0.667.

The attitude In which the landing gear unit is drop tested shall be such as to simulate the airplane landing condition which is critical from the standpoint of energy to be absorbed by the particular unit.

13.355

Reser". energll absorption

drop tests. If compliance with the re-

serve energy absorption condition specified in § 3.352 (b) is demonstrated by free drop tests, the drop height shall be not less than 1.44 times the drop height specified in § 3.353. In simulating wing lift equal to the airplane weight, the units shall be dropped with an effective mass equal to h

W.=W

h+d

where the symbols and other deta1l.8 are the same as in § 3.353. ItEl'RACTING MECHANISJI

General. The landing gear retracting mechanism and supporting structure shall be designed tor the maximum load factors in the fiight conditions when the gear is in the retracted position. It shall also be designed for the combination of friction, Inertia, brake torque, and air loads occurring during retraction at any air speed up to 1.6V." fiaps retracted and any load factors up to those specified for the flaps extended condition, § 3.190. The landing gear and retracting mechanism, including the wheel well doors, shall withstand flight loads with the landing gear extended at any speed up to at least 1.6 V'I fiaps retracted. Positive means shall be provided for the purpose of maintaining the wheels in the extended position. § 3.357 Emergency operation. When other than manual power for the opera. tion of the landing gear is employed, an auxiliary means of extending the landing gear shall be provided. § 3.358 Operation test. Proper functioning of the landing gear retracting mechanism shall be demonstrated by operation tests. § 3.356

~ :.!.;j59 ['osition indicator and warning device. When retractable landing

wheels are used, means shall be provided for indicating to the pilot when the wheels are secured in the extreme positions. In addition, landplanes shall be provided with an aural or equally effective warning device which shall function continuously after the throttle is closed until the gear is down and locked. § 3.359-1 Wheel position indicators (CAA policies which apply to § 3.359). The "means" required by § 3.359 may

tion. In determining the limit airplane inertia load factor n from the free drop

consist of lights of various colors. The signal "all lights out" will be considered by the Administrator as satisfactory If used to indicate intermediate gear positions belt It will not be considered as providing adequate safety if used to indicate either extreme gear lOCked posi.. tioD.

test described above, the following formula shall be used:

[12 F. R. 3437. 36)

§ 3.354

Limit load factor determina-

Correction noted at 14 P. R.

§ 3.360 Control. See § 3.384Where nf=the load !I\ctor developed in the drop tes,t. 1. e., the acceleration (d./dt) in g's recorded in the drop test, plus 1.0.

The value of n so determineci shall not be greater than the limit inertia load factor used in the landing conditions, § 3.243,

WHEELS AND TIRES

Wheels. (a) Main landing gear wheels ' 0:1:.)

Ifl n

D""T

.Tendon old \MC>C.o

apooi- I!oigh'

1/r:uIn

tic

l11l

f/ Clll1 arev1t7

",..te '1' TIl}

(Il)

-:- 10·

lI'f-T-7ge f,w-'1'-llb '!'ub.

110M 11067-II 11oe7.Jit 57-153 57-153

~7-A-I0

l.eo • ISO .ISO .70

Sho.' Shoot Shoot Bar Bar Cotg Cotg C.tg

~?-A.-8

~7-A-I0 f!.6-A.-7

~-277B

~~ 1I-lB&

1.00 1.U 1.35 1.U

~l.e

8.79 2.77 2.77 2.77 2.79 1.79 2.77

Il.es

.101 .100 .100 .100 .101 .101

II

:lilt

114

"730 "610 7._ ,,1111

..~IOO

~"'"

40 1"

'01 I" 110

110

730

....0

7.10

~"'" """'I

i:\

010 010 730 I

1

IE GO

.",

3IlO 380

...,

Spilt'!:.f'IllIt. \ Moduloa m tlUtlcU..

!

1

Num'l t'tDt P.. at

her of c.tI

.

birch

I

I ,

I

PanlloI (1.000 poundl per !qUIrt JIk'b,)I

.., NO "350 ..,~I '."" 1: 1 ..... ...... . .., .....,eo §i I,

100

7._ 7.700 1,140

& ~ I'"1OOM c:.uoEVE.S "" &.

1

""'tVL...TIp\,)(

..... _c

L..OAD$ ..-..oM

~~

ANO l.040

e-t I.S"T

~TIVC&..~

°0

I Z !S .q. 5 '" 7 WIDTH Of" Wooo ME.MBfR. IN INCHE.S

BEA~IIo..IG

STeENGTH OF BeL-TS It-J WOOD - PAIC.ALLE.L TO GeAlN

Fig. A6ba2 5~AI2.ING

t3

~~

~~

~O IF

12-

e ~.."...,.....

A'T ONe. END OfT

eocr 0I0L1')

.q.

I'l£OM

6 ::>

I't>Ir.

ON'OE: 1..000

c~s

"'to
(phi)-Angular deflection. p (rho)-Radius of gyration. !' (mu)--Poisson's ratio. I (prime)-In general denotes an Heffective" or "precise" value.

\,

1.3 COMMONLY ~~i:O FORMuLA;) \,Cmlwi ju'lO-oa) 1.32

SBIPLE eXIT STRESSES

1:1 j/=P/A (tension). 1:2 je=P/A (compression). 1:3 jb=JJy/I=Jf/J:. 1:4 j,=S/A (average direct shear stress). 1:5 j,=SQ/lb (longitudinal or transverse shear stress). 1:6 j.= Ty/l p (shear stress in round tubes due to torsion). 1:7 .1,= T/2At (shear stress due to torsion in thin-walled structures of closed section. Note that A is the area endosed by the median line of the section). 1.33

CO~IBINED

STRESSES (see sec. l.535)

1:8 jn=f,+fb (compression and bending). 1:9 j'mu= "{F.+ (jn/2)2 (compression, bend-

ing, and torsion). 1:10 j"max=j,,/2+j'max' 1.34

DEFLECTIONS (Axial)

1 :11 1 :12

e=o/L (unit deformation or strain). E=j/e (this equation applies when E is to be found from tests in which f and e are measured). o=eL=j/EL =PL/AE (this equation applies when the deflection is to be calculated using a known value of E).

1 :13

1.35

DEFLECTIONS (Bending)

1 :14

di/dx=M/EI (change of slope per unit length of beam, radians per unit length).

1 :15

iz=il+ (rt M/EI dx=slope at point 2.

Jrl

(The integral denotes the area under the curve of M/EI plotted against x, between the limits Xl and X2.) 1 :16

Y2=YI+iI(X2-XI)+

JIIrx'JJ/EI(Xz-x)dx

=deflection at point 2. (The integral denotes the area under a' curve having ordinates equal to J;J/EI multiplied by the corresponding distances to point 2, plotted against x, between the limits XI and X2.)

1 :16a

A6b:ll

.'12=YI+ [I'idx=deflection at point 2.

JXI

(The integral denotes the area under the curve of (i) plotted against x, between the limits Xl and .f2). 1.36 DEFLECTIONS (Torsion) 1 :17 d¢>/dx= T/GJ (change of angular deflection or twist per unit length of member, radians per unit length).

1:18 c/>= (XI T/GJ dx=total twist over a JXI length from XI to Xz. (The integral denotes the area under the curve of T/GJ plotted against x, between the limits XI and xz.) 1 :18a ¢>= TL/GJ (used when torque T is constant over length L). l.37 TRANSVERSE DEFOR~IATIONS 1 '19 _ / = unit lateral deformation . J.I.-eL e unit axial deformation (Poisson's ratio) 1 :20 Ee x ---}x - J.Liv. 1:21 Eey=jv-JJ:ix. l.38 BASIC COLU~IN FOR~lULAS 1 :22 Fc o=c7r z E/(L/ p)2 (Euler formula for long columns). z =7r E/(L'/p)2 where L'=L/..r;. 1 :22a Fc=c7r z E' /(L/p)2 (modified Euler formula for short columns). 1:23 Fc=Feo[ l-K(L' /p/7r.J E/Fco)"] (general parabolic formula). 1 :24 Fe=F,"~: - Fcu(!.'/p)2/47r 2EJ (2.0 parabola-Johnson formula). 1 :25 Fe= Feo [1- 0.3027( eL' / p) /1r.J E/ Feo )1.5] (l.5 parabola). 1:26

F c=Feo [I-0.385(L'jp)/1r.JE/Fco] (1.0 parabola-straight line formula).

BASIC COLUMN FORMULAS (Nondimensional) 1 :27 R.=Fc/Fco (allowable stress ratio). 1:28 B=(L'jp)j1r.JE/Fco (slenderness ratio factor).

1.39

1 :29 1 :30

Ra= (1!B)2 (E'Jlef formula).

Ra = 1- KEn (general parabolic formula).

Ra= 1-0.25B2 (2.0 parabola-Johnson formula). 1 :32 Ra= 1-0.3027 BI,5 (l.5 parabola). 1 :33 Ra= 1-0.385E (l.0 parabola-straight line formula). 1 :31

STRENGTH DATA

A6b:12

Fonnulas for Stress and Strength, 'Steel and Dural Sections Ra = 1 - x:e h ~O (General p&rabolio formula)

~o

2

Sl

Ra = 1 -

32

R.,

=1 _

3~

~

= 1 - .S85B

of LocdblG

.25B

(2.0 parabola - Johnaon formula)

.302781 • 5

(1.0 parabola - .trai~t line formula)

Unit Stress

SHEl.R LlID COill'F~SS!OlT

f

~~-1 ,L.P

c

P Dt

:

• V

f

ot

tI

~J~V

l

(1.5 par&bola)

l.llom'.b1e Unit Strt:ss I fa' =jfs2 + (fj2)2 not to e;~ce,;d p.llol'mh Ie for pure shear fo' : f0l2 + fa' not to exceed E--for compression.

Refere:rc~

POOrIJC'.n P. 215

Fle.t plate

I

SHEI-R IJiD

t

p

.

f t • Dt

TrS~ON•.

'p

tilr:l~lV . Flnt

f

- V

8

ot

-

s

t

not to exceed

1 f t ' : ft!'2 + f s' I

Ft for tons ion•

Plr-to

I B:l!:Dnm loHD S:IE'J..Tl

,-

1'0

'O?ti~ IIf

I

"'iok syr.motrlo~l

,

, Uns~TIn(trio('.1 soc io:-. londo( nt elas'.::ic axi

::-.bClut y r.xis

~ 1--:'1

f

=!:fl• SQ

l'S

s

I

fo=¥y-

~

fs •

Thb box

f

Type of LoB.di~ Bending and Compression Sprucs box or I section

t;j

I

Round tubes of Steel or ~re.l

0

!

II

r

fc

=stress in thin

= SQ

fs

= r.llOTI'l',ble

Sib Unit Strese

fb: fc

(500 clso TIT 5~3) Fc;.for Wt.:c;120 :.3 ./fu'co (t/~) I for 1:, t >-120

• ):., C

sec·~io:.l

not to e::oel!.d E'.llm7l'.ble o for conprossion

f s not to exceed allm'/F.ble for puro shear

,--

Thin oyl1r..der

f

-

POOrtlflll p. 215

i

~'"

I

2 + (rcl2) not to 8 r.llowmble for pure shear

, :/;

e7.oewd

=

¥

=

511.1

~11~b1e Qonpr()s~ive f1~t plato

Gher.r stro81S

fer thin flr.t plnte

AIlowable Unit Stress

Reference

ANC=/If>

See P A3:6.1

Sea. 2.41

t

f b -!:!r - I P fo A

Eq.

fb

1);

fc

~ Fc

ANC#5

= 1.0

Fb : bending Mod. of rupt. Fc : allowable oolumn stress

.

,~-

Seo.

4.40

~

See pp. 5 to p • -6. v M - bending moment about neutral axis, in. Ibs. y _ distance from Jlelltr'-ll axis to o;l1;.or f 1bprs. I = moment of inertia of area aoout n~utral axis. S - shear, Ibs. Q - A'y and b width at point of shear stress

Notat~on.

IT

=

A'

Ux;i t str~:'!!l

TYPe of Loc.dh£

~@

t s :-,-

,•_

~

Tbin lim

.r:-·: -,""

_

cy

t

2T

s .. ;rcI2l:

Lh ~/2

1T

~

=

ill.e1l. TN

~~~~.~s n~b 6.t" E(t7b)~

~~

Aoro. Oct. 133 p. 1-:'7

H~:~ E'f~lh2r*-b 0 \ 1 1

:Tith ~~>Jdt 6.5.JEFcdt/n)

p/J.

SI

.L

,0."

4. 5 (b!l:'. + .J08) E (t/b)2 but net narc -!;han Ft/2 '. j. Diq;o~.l tensio:l yield at F 2 tt (-i{!'.t;.lor "tund":l fiuld" sh(J('.r brnoinc)

III

CD

:;" Eq. 5&?'.1

=--=..-=~~====

v/bt ::

I

Tiruoshanko

'1

~

'1 CD CD

til

II>

c-)

:s 'J Po 1 Vpo~~fI en •: •

i

.... '1 " H CD

::t~

~~ ::r't-;!

.. > ~ CD

(For corrugated, see A3:l3)

I!-

Notation:

8. ~

L' - length between bulkheads, inches. t - plate or sheet thickness, inches. d cylinder diameter, inches. L - length or column, inches. ~= least radius of gyration of section, in. E - modulus of elasticity in compression, lbs./sq.in. Ft;v- tension test yield point, lbs./sq.in. Ftu - tension test ultimate strength. P - load, lbs. A - area of cross section, sq. in. .st'~ Q/S'ci

=

'" Notation: {See also p. P:3-4) L length between bulkheads, in. A - area enclosed by boundary F shear stress at yield; Fty • tensile stress at yield; Ftu ultimate strength (tension). r~J· thiCkness, in. d - diameter. in. J . 1rr4 for solid shafts. J - 11' L!'"f' - (r-t) 4] /2 tor tubes. 2 outside radius of rod or tube, in. T ;: twisting moment, in. lbs. r

=

u

~

Thbnp

f.vidion Hnndboo!:: p. 1.,30

8

Ft

~~\rl~t~h~-dl/~1t_>~12-~Q~--~~-=~--......a .. .a .. . . . Aviation SHE.t.R F1 t Eucl::liDj; fr.ilur\l P.t Hr.ndbool: po ~30

"'OF

21

J.}lD TENSION

~

6E(t/b)2

T

:=.

-

.

~po 130

I

Refer"n~

Fty Cot yiold pcint

p/l.

PInto ~d-es su J)c.rti..d p/J_ (fo p"-"d" : .. 1 ) or oorruba~ G,e p.~1

~~~~~~~---~~-~-----:----L-------------_1-----_l~ b

J.J.lo\"1r.blo Un! t streGs

~Q,~ect;uns '" ?E/(L/r)2 tor ?/J.. 10

/

/

r-

.......

/

/

/

/

~

V

I-"> i-"

V-

i-"

V I-- I-- ~

V

20

30

40

50

91 FIGURE 2.42 (b).-ToJ'8ional modulus of rupture of round alloy steel tubing.

60

A6b:l7

(ReprodUCed wi th permission from .4NC-5a) TABLE 3.111 (a).-Duign 17I$chanical propertu.a oj bar. t4S shut and pIaU (kips per square inch)

I

st.tand plale'

AN-A-12

:=.·.···············.·········.···.····.·b CoD.d.JUaQ_. _____________

~

~.

__________________ • ___ _

•••• _. ____ .. _____________ ._. ____ • _____ .__

1"..: than l.S l)lu .. t ht· "ul, .. tanti.ateri hy srif'qualf' te~t:- pilar to apprt,,'all,.\ till' l,rO('llrlll,Q: or certificAlilJp;

AJi!:ell('\,.

98 124 ilO

91

0.101

_ ... __________

Hardclad R30I-T42

ConmH:'r('ial desig- : lIalioll ..• _

H&rdclad

I

R301-T3 A1cl.d

I TABLE

O..m-

--:-:--:-1---:---' :---.:----:---:---:- ~-:----:- "-6~ i

1':. R.. G. WIIl .. in.l..

O.UiI-0.4519

I-A--!-A-,-A-;---I--·-.,-!--'-A--- - - - - - -

T .... : F.,......

Hrat-treatt'd

0.2.50- i O.m- 1fer~ to all material ~l1ppli{'rl. ill the annealed temper and heat-treated by th(> u"er, a.lld to"all material re-il(>at-treated In' th£' lI~er reKardle ...;,.; of the temper in which the material wa ... :'llpplied. . b For (>~tr\l"iOf:"; with ()!]t~tan(r",_' I"'T": '111::S ~. Si1:3! ,,!:. I • ~ 8~ ~~

1---1--I~-I~-

~~-::-----i~, 24S-T4 ----I A51S-T6________ 61S-T6..___ 75S-T6_______

40! 32

I

25 I 20 291 21 115.5 12. 5 31 I 22. 5 17.5 15 451 35 27 22

i

I i I

18 11 13.5 21

I Tests conducted by Aluminum Co. of America on specimens up to 2.O-inch diameter show no appreciable size effect.

3.112 (b).-Repeated-j/eX'Ure jatigtu 8treng/JI oj aluminum aUoy 8heet 1

T-'BLE

Data based on testa conducted by Aluminum Co. of AmeriCL TABLE

S.

C},

±331 ±331 ±32 ±32 ±31, ±291 ±271 ± 25 ±26 ±27 ±26 ±25 ±23 ±21 ±191 ±38 ±37 I ±36, ±34 ±33 ±31 ±29

-5

~1.

3.112 (d).-Shear jatifl!U strength

1

[Values given were determined by testing O.330-inch diameter specimens in Aluminum Research Laboratory torsion fatigue machine]

[Values given were determined by testing O.OO4-inch thick sheet specimens in Aluminum Research Labora.tory repeated flexure fatigue machines, and represent the completely reversed stresses that such specimens will withstand.]

ReW'ned (alWnAtinJ) !tn!sII, ta. at indlcsted number 01 cycle! (steady)

I

"''" I

14S-T6 and 24S-T4 ____________________

75S-T6 _______________________________

I 1

0 +5 +10 +15 +20 +25 +30 0 +5 +10 +15 +20 +25 +30

100,000

",cles

1,000,(0)

I

~

±27 ±26 ±25 ±23 ±22 ±2O ±19 ±33 ±31 ±3O ±28 ±26 ±24 ±22

Data bued on tests conducted by AlumlDum Co. of America.

±21 ±2O ±19 ±18 ±17 ±16 ±15 ±23 ±22 ±21 ±2O ±18 ±17 16 ±

1

10,000,(0)

"',. ±16 ±15 ±14 ±14 I ±13 I ±12 ±11 ±19 ±18 ±17 ±16 ±15 ±14 ± ,31

100,000,000

m'"

±ll ±ll ±ll ±!1 ±IO ±IO ±9 ±16 ±15 ±14 ±14 ±13 ±12 ±!11

"',.

Fatlrue strength (reversed stJ't':Slll) psi at. indicated number ot cycles

500,000,000

±I o ",1 o ±I o ±9 ±9 ±9 ±9 ±I 4 ±I 3 ±I 3 ±I 2 ±I 1 ±I ±I-

i

:1-

~.lI § §! h §'" §" - - - - - - - - -- -- - - - --0

Alloy and temper

~!

1

~t'

-

Alclad 14S-T3 ___ 31 Alclad 14S-T6 ___ 31 24S-T3 __________ 34 24S-T36 _________ 36 Alclad 24S-T3 ___ 32 Alclad 24S-T36 __ 32 Alclad 24S-T8L _ 27.5 Alclad 24S-T86 __ 32 61S-T6 __________ 29 75S-T6 __________ -----Alclad 75S-T6 ___ 29

I

!

0

20 20 27 27 20 20.5 17 19 23

25 20

g.~

8", ~Q

17 17 21 21 15 16 14.5 14.5 17 21. 5 15

15. 5 15.5 18. 5 19 13 13. 5 13.5 13. 5 13 21 13

15 15 17.5 18 125 13. 5 13. 5 13 12 20.5 12. 5

I Data based on tests conducted by Alummum Co. o! America.

STRENGI'H OF

.

~

r

P... RT

"

• '. -

-~,

ACTUATORS

-.... -TABLE ·OF STANDARD MODELS #,

. '., .

_- - _ _ -

BA~SCREW

..

Y

~

~

"-.

_ - ,

NO.

PfTCH 01 ....

NO. OF TURNS

lEAD

5401 5402 5403 5404 5405 5406 5407 5408 5409 5410 5411 5412 5413 5414

.500 .625 .750 .875 1.000 1.250 1.500 1.750 2.000 2.500 3.000 4.000 5.000 6.000

6.5 7.5 8.5 10.5 10.5 10.5 10.5 10.5 10.5 10.5 10.5 10.5 10.5 10.5

.166 .166 .182

.093 .093 .109

.~vv

.·12~

.200 .250 .286 .308 .333 .364 .400 .500 .572 .666

.125 .156 .187 .218 .250 .281 .312 .375 .437 .500

B... LL 01 ....

NUT O. O.

.750 .937 1.062

1.250 q75 1.687 2.000 2.312 2.656 3.250 3.812 4.968 6.125 7.312

LENGTH

R... D. OVER TUBE OR ClAMP

"0"

1.875 2.125 2.562 3.187 3.437 4.062 4.625 4.937 5.375 5.937 6.437 7.687 8.687 9.937

.812 .875 .937 1.062 1.125 1.375 1.500 1.687 2.062 2.312 2.625 3.437 4.125 4.812

1.812 2.000 2.500 2.312 3.125 3.625 4.375 5.125 5.250 5.875 7.000 9.000 10.500 12.125

NUT

'_

SCREW BORE

'T'

MAX.

.125 .125 .187 .187 .312 .312 .375 .375 .437 .500 .562 .625 .750 .875

.257 .383 .468 .550 .675 .843 1.012 1.187 1.343 1.843 2.187 3.000 3.860 4.703

.. 00 300 200 SCREW No.

I

V>

0 Z

'"2 0

100 90 80 70

'"0z

'0."

30

;:'i ~

I

5412

60

50 .. 0

5411 5410 5409

20

0

5408

9

5407

0(

10 9 8 7

6

5406

'.1

100

1000

UFE IN MlLUONS OF TOT... L INCHES TRAVEL

Ball-screw actuator data.

~

Stock sizes with screw leng!hS frolY 2 inches to 20 feet, • ~_ T:hese $ta,!d:

~

Lightplane ..

Plane A ••.

58 17

----rw IT. 40 I': Q)

I':

~

20 0

"':-..~~ I'"

110"" r-......

I--

~ :~ r--... !--.. ~ ....... 1:' ['('~.~

,

"

100- I""- ~ ~ r-.... ""'""~ I-- ~~ 80 Ii !I 70 --. ~ 60 -r-.50

40 30

......~

~ :....~ ~~ ~

Ventilating system: l/Sinch felted kapok covered with pliofilm. Cabin: Two l-inch dry z., blankets next to hull, various thicknesses of felted kapok, casement cloth. Materia.l supported by hooks which ere riveted to fuselage. Seapak is attached to hull with latex cement and may be reinforced by metal strips. 1.lateria.L is glued into place with Dum Dum.

Cabins and Cockpits: Seapek and felt cemented to metal surface with Vultex cement. Rubatex attached directly to sides and deck, covered with neoprene cement. Sprayed direct17 to akin. Sprayed or brushed on rubber. Sprayed or brushed on metal - not on felt. Ventilating ductal line with lie-inch felt.

'4 ~1-

120

:t

t:--

Location and method ot application Cabin: Insulite and Seapac nailed to wooden cabin framing.

,.

Fee

r-.::j -§t-

I

20

t

-..... -

-

0-- r----

100 Airplane Motor 90 N.I.SUbW81,50 mph

./L

~4- 60 Auto,con9rete ./'

rr

80 Pullman sleeper

...

3::..

-r-

10':--

I"-

.J-

K', ./- -

highway.4U mph

40 Usual

conv9rs~tion

20 Sort music

T [,,- o Rustline leaves t

t

jb above 1 mb. I 100 500 !JOO 500b 10000 Frequency in cycles per seconi Fig. A8.4 Lou:1ness level contours of the ear; each contour represents all the tones which a.re equally as lout a. a 1000 c,vcle per second note. For example, an 80 cycle per ttecod note at 7(\ iecibeh sounl. as loui a. a 1000 c 'cle- note at 50 iecibeb. (N.A.C.A. TN' 748) " t

948)

I

..... r-......!-.

--1

, 1

,I

NOISE DATA

A8:3

1\

\0

o

la, Ib,

,, I

c \\ i\ \

Ie,

IIa, IIb,

,, 'l\na\

\

~ ~

are not noticeable just noticeable well noticeable very strongly noticeable disagreeable very disagreeable

\

UbI

"\

Ie

'Ib\ n IJ 1\ B

01\' 1\

r\ \ \

:LS.:

~ '\

,A '\..

~

3

"' ~f\ "' 10~.1

1020 466OFig. A815 RelPOI1H or thl 11141T1411&1 Frequency, c p B to .,lbratioJl (N.A.C.A. TN 748)

1

tlirst attempt to soundproof a~ airplane ,l'iei?,ht of soundproofing p"r passenger , /' / /NQise lev"l in db. , "'! I 30 It I < , : , _ /. .... Curt~s "S.:lpi)r" ,!,-32 I "

,