An Experimental Investigation of the Effect of Propellers Used as Aerodynamic Brakes on Stability and Control

FI LE copy NO . I- W

NATIONAL ADVISORY COMMITTEE FOR AERONAUTIC S

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ORIGINAll Y ISSUED July 1945 as Advance Restricted Report 5COl AN EXPERIMENTAL INVESTIGATION OF THE EFFECT OF PROPELLERS USED AS AERODYNAMIC BRAKES ON STABILITY AND CONTROL By Victor 1. Stevens, George B. MCCullo ~ u :;. gh ~ ' ~ ~-t'aq~t:Jrvl and Frederick H. Hanson ".

~ be ,etill" d to -

thtrl11M of 'tOO !' "a tiOilm -

Ames Aeronautical Laboratory ~sory Comn d etr Moffett Field, California

- tar ~el onal:lt i cs

,Ytasbingum, .0. C.

NAClA

W ASHINGTON . , [ I NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassliied. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution.

W'f A -19

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. ~ NACA ARR No. 5COl NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE RESTRICTED REPORT AN EXPERIMENTAL INVESTIGATION OF THE EFFECT ON PROPELLERS USED AS AERODYNAMIC BRAKES ON STABILITY AND CONTROL By Victor I. Stevens, George B. McCullough, and Frederick H. Hanson SUMMARY Tests were made of a model representative of a single- engine tractor-type airplane . for the purpose of det e rmining the stability and control effect s of a propeller used as an aerod J' namic brake. The tests were made with single- and du al- rota t ion propellers to show the effect of typ e of propeller rotation, and with positive thrust to provide basic data with which to compare the effects of negative thrust. Four con- f iBurations of the model were used to give the effects of tilting the propeller thrust axis down 50, raising t~e hori- zontal tail, and combining both tilt and raised tail . Re - sults of the test s are reported herein .

The effects of negative thrust were found to be signifi- cant. The longitudinal stability was increased because of th e loss of wing lift and increase of the angle of attack of the tail. Directional stability and both longitudinal and directional control were decreased because of the reduced ve- locity at the tail. These effects are moderat e for moderate braking but become pronounced with full-power braking, par- ticularly at high values of lift coefficient.

The effects of model configuration changes were small when compared with the over-all effects of negative-thrust operation; however, improved stability and control character- istics were exhibited by the model with the tilted thrust axis. Raising the horizontal tail improved the longitudinal

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NACA ARR No . 5COl characteristics, but was detrimental to directional character - istics . The use of a dual-rotation propeller reduced t4e di - rectional trim charges resulting from the braking operation.

A prototype airplane was assumed and handling qualities were computed and analyzed for normal (positive thrust) and braking ope ration with full and partial power. The results of these analyses are presented for the longitudinal characteris - tics in steady and accelerat ed flight, and for the dir e ctional characteristics in high- and low-speed flight. It was found that by limiting the power output of the engine (assumi~~ the constant-speed propeller will function in the range of blade angles required for negative thrust) the stability and control characteristics may be held within the ltmits required for safe operation. Braking with full po wer, particularly at 1m. speeds, is dangerous, but braking with very small power output is satis - factory ~rom the standpoint of control . The amount of braking produced with zero power outp ut is equal to or better than that produced by conventional s~oller-type brakes.

INTRODUCTION Modern aerial combat experience has demonstrated the need of a device to produce rapid decel eration of certain tactical - type aircraft for the foHm"ing purposes: (1) to allow more time for aiming and shootin g at a slOw-moving target after it has been overtaken; (2) to reduce the time required tO,slow dmm for torpedo launching; and (3) to limit the maximum speed in a dive . Although not covered in this report, a f ourth pos - sible use of a powerful decele rating devic e is to shorten the land ing run of an a irp1a.ne after.. it has made contac .t with the ground . .

Previous tests of a model equipped with flap - or spoiler- type aerodynamic brakes have shown this type to be subjec~ to s~veral undesirable features: ( 1) change of trim of the air- plane resulting from application of the brakes, thus spoiling the pilotte aim on the target; (2) t ail buffeting resulting from the turbulent wake of the drag - producing flaps impinging on the tail} (3) loss of effectiveness "'ith decreasing speed , and (4) increased complications of the wing structure and dif - ficulties of prod u cing a smooth exterior surface on the wing.

- --~- . ------ -- NACA ARB No. 5COl Wind-tunnel tests have been made to determine the effect- tivemess of a propeller as a brake (reference 1). It vias found that negative thrust could be more effective in slowing down the airplane than the increased drag caased by flap-type brakes. Curves showing the computed variations of airplane speed with time are shown in figure 1 for both types of brak - ing. No measurements were made, however, of the effects of negative thrust of a pro p eller on the stability and control characteristics of the model .

It is the pur p ose of the tests reported herein to deter- mine the ef f ects of a br a king propeller, as it might be used in actual serVice , on the stability and control characteris - tics of a wind-tunn el model. The tests were made in the Ames 7- by lO-foot wind tunnel . The model used for the tests is not a soale model of a particular airplane, but rather is representative of a general type of highly loaded, highly pow- ered, single - engine) tractor-type military aircraft . The s cope of the tests was intended to be s u ffici e nt to cover th e use of the propeller as a brake in level and diving flight .

No tests to simulate the landing condition were made.

Some brief preliminary tests made with another model in- dicated that the use of negative thrust produced significant changes in the stability and control characteristics. For this reason the scope of the present test was expanded to in - clude configuration changes which it was hoped would result in improved characteristic s with negative thrust , and also to provide a basis for estimating the characteristics of air- planes of different conformations. The configuration changes include tilting the propeller thrust aXis, ra1sing the hori - zontal tail , and a comoination of both . The model in all configurations was tested with both single - and dual-rotation propellers .

MODEL The model used is representative of a midwing single - en~ine airplane of a type requiring additional means of speed control because of its tactical purpose . The fuselage lines have been simplified and do not include a canopy . For con- venience, the model is referred to as the stability model.

A three-vi ew drawing is given in figure 2 and complete

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4 NACA ARR No. 5COl dimensions ln table 1. Photographs of the model installed in the wind tunnel are given in figlITe 3 .

The horlzontal tail could be mounted in two positions : 4.26 inches (basic configuration), and 1.2. 77 inches above the fuselage reference line. Also, · the thrust axis of the pro - peller, which normally coincided with the fuselage reference line, could be inclined in a vertical plane about a point ap- proximately midway between the two propeller disks . A sketch showing the relation of the propeller and horizontal tail to tho center of gravity 1s given in figure 4. A key to the con- figuration notation used is give in the appendix .

Both the · single- and dual - rotation propellers were mountod in the dual - rotation spinner. The front hub was keyed to the motor shaft) and drove the rear hub through reversing gears.

Four blades were used in both the single- and dual-rotation pro - pellers . For single rotation, four blades were placed in the rear hUb . Since the rear hub will accept left - hand blades only loft - hand rotation was used for all tests with the single - rotation propeller . The propeller diameter was 2 . 52 feet .

The blades used were models of Hamilton Standard blade forms Nos . 3155-6 and 3156-6.

A prototype airplane was assumed in order that the wind - tunnel data could be applied in terms of handling qualities of a full - scale airplane. Tho dimensions of the prototype were such that the model tested became a 3/16 - scale replica of an airplane possessing the following characteristics: (1) weight,

14,700 pounds (W/S = 39 . 2 lb/sq ft), and (2) power J 2100 brake

horsepower at 1350 rpm of propell e r.

POWER CONDITIONS In order to convert the wind-tunnel data into airplane operating conditions, relationships between thrust coefficient Tc and lift coefficient CL are required for the various c onstant-power outputs to be considered. To achieve a constant - power output with negative thrust, the propeller-bLade-angle actuating mechanism must be capable of functioning in the neg - ative blade-angle range in the same manner as in the normal positive blade - angle range for positive thrust. To avoid overspeeding the ongi~e, the blade-angle actuating mechanism [ NACA ARR No. 5CQl must be ca p ab l e of rapid motion through thG windmil l ing range .

It may be well to review the operational states of a propeller as the blade angle is changed from positive to negative , Assume for simplicity that the propeller operates at a constant value of V/nD (constant speed of the airplane and constant rpm) . At a large positive blado angle, the propell er produces positive thrust and absorbs power from the engine .

As the blade angle is progressively reduced , the propeller ab- sorbs less power and produces less thrust until it no longer absorbs power and is furnishing a small amount of negativ e thrust because of its own rotational losses . This is the be- ginning of the wiridmilling state , and the propeller now tends to driv e the . engine and will cause it to overspeed unless a brake is used to hold the revolutions per minute constant.

The n ega tive thrust produced is progressively increasing. A blade angle, however, will be reached at which the propeller no longer tends to drive the engine . This is th e zero-power condition and , is accompanied by a substantial amount of nega- tive thrust. Further reduction of the blade angle will CaUse the propeller to windmill backward unless power is supplied by the engine to keep it turning in the normal direction.

This is tho beginning of the power-on negative-thrust state; the amo lm t of negative thrust may be increased by increasing the e ngine power.

The relationships of Tc to CL (fig . 5) were computed fr om the data of reference 1. A rate pm·Ter of 2100 brake hors o power at a propeller sp ee d of 1350 rpm and a wing load- ing of 39.2 pounds per square foot we re assumed. Curves for zero power at 1350 and 1000 rpm are also shown in figure 5.

It will be noted that, for low values of CL (corresponding to high values of V/nD), only a small increase of negative thrust results from the use of full power. An infinite num- ber of families ' of Tc versus CL curves may be obtained by varying the power and r ev olutions per minute, and it should be possible to find a power condition which will sat- isfy the operational requirem e nts for negative thrust, TESTS The wind-tunnel tests were made of the model in th e fol- lowing configurations: (1) Model in basic configuration

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6 NACA ARR No, 5COl (2 ) Model in basic configuration with th r ust axis tilted dovm (3) Model with raised horizontal tail (4) Model with raised horizontal tail, thrust axis tilted down 50 Tests were also made with the tail r e mov e ~.

In addition to tests with negative thrust , tests were mD.QC with the propeller removed and with positiv e thrust to serve as bases of comparison for th e effects of negative thrust.

Te sts made with the single-rotation propeller were duplicated with tho dual - rotation propeller to give the effect of type of propeller rotation. All tests with power were made at constant thrust .

Some preliminary negative - thrust tests were made with propeller - blade angles of -50, - 100, and -150 (measured at the 0 . 75 radius station). The effect of blade angle on . the stability characteristics of the model proved to be inappreci- able , and J since a blade angle of -150 gave the · be · st condi - tion s for tunnel operation, all further negative-thrust tests were made with this blade setting. Tests with positive thrust Were made with a blade angle of 25° , Experimentally determined Tc versus V/nD relations for these blade angles are shown in figure 6, Longitudinal Tests To determine longitudinal - stability and longi tud.inal- control characteristics of the model) tests in pitch were made of the model in all four configurations with various elevator deflections .

Directional Tests Tests in yaw to provide directional-s t ability and di r ectional - control characteristics were made of the model in the basic configuration and with the thrust axis tilted down 5° with various rudder deflections . The model with raised horizontal tail, thrust axis untilted and tilted) was tested

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---- -.---~~~- NACA ARRNo . 5COI 7 The tests in yaw were made with the rudder undeflected only .

0 0 at two angles of attack , ' o. = _2 and au = 9 .

u COEFFICIENTS AND CORRECTIONS The data are presented. 'in NACA standard coefficient form and are corre ' cted for tunnel - wall effects. The corrections were applied to the negative - thrust data in the same manner as for positive thrust because of lack of information on tunnel - wall effects with negative thrust . No corrections were ap - plied for strut - tare and interference effects. Previous ex- perience 'Vlith s imilar models has shown that the corrections are small and have no appreciable influence on stability and control characteristics . The dimensions on which the coeffi- cients are based and the tunnel - wall corrections applied are given in the appendix.

Moment coefficient s were computed for a center of gravity located fore and aft by the 25 - percent point of the mean aero- dynamic chord and I-percent mean aerodynamic chord vertically a'bove the fuselage reference line .

RESULTS Tests to Determine Longitudinal CharacterIstics The large number of figures involved makes it impractical to present all of the constant - thrust data for the four model conflguratlons tested. Instead, complete constant-thrust data for the basic configuration of the model, single - rotation pro- peller", ar e presented. , However, longitudinal characteristics corresponding to various pO>ler conditions obtained by cross- plotting the fundamental data are given for all model con- figurations .

The wind-tunnel data obtained with the model in the basic configuration, single - rotation propeller with positive and.

negative thrust, are given in figures 7 to 13 . Data obtained with the tail removed are given in figures 14 and 15. Data obtained with the propeller removed, tail on and off , are given in figure 16.

NACA ARB No . 5eOl Longitudinal characteris ti cs for the four model configu - rations are presented in figures 17 to 32 . For each cOD~igu ration) rated power was simulated with single- and dual - rotation propellers for positive and negative thrust. Summary plots of the variations of C and C with CL to sho1-1 m h e the effects of model configuration with either single - or dual-rotation propellers are presented in figures 33 to 36 .

The effects of a change in incidence of the horizontal tail for the ba s ic ccnfiguration are given in figure 37 .

In order to shoYl the of fect of pO'ver on the longi tu d inal - stability characteristics of the model with basic configlrra - tion, a summary :plot for various power conditions is given in figure 38, To investigate the effects of the sli:pstr eam on the ho:dzontal tail) velocity surveys "ere made in a vertical plane con t aining the . elevator hi.nge line. The data, plott ed in the <ltail form of contours of equal values of the ratio qfree stream are presented in the fol101ving figures:

Fig . no . I Configuration a.u

(deg) SP - _HV 39(a) - 0. 1

o

S SP " -1 5 _ h'V 39 ( b)

o - .2

o SP - 15 _ HV

40(a ) 9 -. 1

s

SP - 15 _

40(b) HV 9 - .3

S SPS - 15 _

40(c) HV 9

- .5 For the :purpose of Gaining further insight into tho sta - bility and control effects of negative thrust) the ratio NACA ARR No , 5COl 9. e ffective at the tail anQ the angle of attack at the tail 9.free stream ot were computed as follows :

9. t = (dCmld~e)~ u power on

9. (dCmld8e)~p ower off (subscript t Q enote s tail) These results for rateQ-pOI'ler ope ration are compared in figur o 41 for the four model configurations.

The computed. variations of elevator angle and stick force with indicated airspeeQ in steady flight are presented in fig- ure 42 for all four model configurations, rated -p ower operat1on.

The longitudinal characteristics corresponding to zero pOlver at a propeller speed. of 1000 rpm for the model in the basic configuration, single-rotation propeller, are presented in figure 43. The longituQinal characteristics with elevator deflec t ed for the other config l~at ions of the model will not be presented, because almost iden t ical elevator effectiveness anQ hinge-moment characteristics were found . However, longi- tuclinal characteristics for all model configurations with the elevator undeflected are presented in figures 44 and 45.

Computed variations of elevator angle and stick force with indicated airspeed in steady flight proved to be nearly identical for all model configurations with this power condi- tion , and are presented for the basic configuration of the model only in figure 46 .

The computed variation of elevator angle and stick force vlith normal acceleration in a dive pull-out (Tc = -0.13, TAS

= 310 mph) is pres ente d in figure 47. To serve as a basis of

comparis on , the variation of elevator angle and s tick force with normal accelera tion in steady turning flight for posi tive - thru st operation is presented in the same figure . Since all configurations gave similar stick-force gradients, results obtained with the model in ba sic config uration only are presented .

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10 NACA ARR No. 5eOl Teste to Determine Directional and Lateral Characteristics

The tests at C1u = _2 with positive thrust were made at

Tc = 0,03 to give characte ri stics corresponding to high-opecd

flig ht with rated pOHer and with negative thrust at Tc = - 0.19

J to giv e characteristics corresponding to diving flight with rated -p ower braking, . Resul ts obtained : \'li th the rudder deflected are presented in figure s 48 to 52 .

Tests at c:. = 9 with positive thrust were made at Tc u = 0.32 to g ive characteristics corre~rponding to climbing fl i ght .lith rated p ow er. To obtain charac ter istics corre sp onding to slm,, -sp eed decelerating flight ) tests with negative thrust were

made at Tc =- 0.19 and Tc = - 0 . 38 . The maximum negative

thrust available with rated. power corresponds approximately to

Tc = - 0.38, and Tc = 0.19 was s ele c te d arbitrarily as corre -

sponding i"o one-half the aV9.ilable negative thrust.

Characteristics of tho model with rudder deflected, single- rota t ion prop e ller, are presented in figures 53 to 59 . The r l'. dder effectiveness obtained with the dual-rotation prop e ller was simi la r.

For the purpos e of shmving the effe c ts of changes in mo del configu~ation and type of propeller rotation on the directional- s tab ility characteristics, results of tests with the rud . der 0 0

undeflected au = _2 and. au = 9 , are presented in the

J ,summary plots of figuxes 60 to 6° , Comparisons of the effectiveness of the rudder are made in figures 70 (a) and (b)J in which the ratio (dCn/o.Or)po.,er on (d Cn/dOr)power off ' is given for the basic mode l co nfiguration .

The effect of power on the lateral characteris t ics of the mo de l, basic configuration, is presen te d for the high- sp ee d condition in 7 igure 7 1. Since the effects are small the variation of Cz with W only is presented. For the low- speed condi tion the e ffe cts are conSiderable, and the variations of CL, Cm , Cy , and C7, with Ware _________ .....J N~CA ARR No , 5COl presente d in figures 72 an d 73 for the model with the tail on and off . To give an in d ication of the effect of power on lateral- d irectiona l correspondence, the ratio of dCn/d \II to dCI/d\ll is plotted a s a ftUlction of Tc in figure 74.

DISCUSSION ·Longitudinal . Characteristics To facilitate p re s entation and analysis the longitudi - nal characteristics are discussed under the headings: Steady flignt with full power, Steady flight with partial power, and Accelerated flight . liandling requirements for a fighter or torpedo bomber are used as a basis for judgment of satisfac- t ory steady-flight characteristics . Elevator contr ol in a dive pull - out would be critical for a dive bomber and is therefore l ~ sed as a basi s for judgment of satisfactory accelerated - flight characteristics. Beca ). se of the similarity of results obtained with single- and dual-rotation propellers, all tUlquali f ied statemen t s apply to the model equipped with either propeller type operating at negative tl1rust .

It will be noted tlllit a · large part of this discussion is devoted to comments on the handling qualities of this model.

This is considered justifiable since it is typical of tl1e existing hig111y powered, single - engine airplanes and its char- acteristi cs are probably representative. On the other hand these characteristics may be altered by either a movement of center of gravity (which would translate the moment curves and cause different sections of the curves to be utilized for trim) or by a ctange in tail plan form (which would c ause the tail to respond differently to the influence of qt and ~t) .

For this reason, due regard should be given to qtJ~' and the tail - off characteristics which indicate the more fundamen- tal effects of a propeller producing negative thrust.

steady flight with full power. ~ Application of full power to a braking propelier produces very high deceleration by re- ducing the air velocity through the propeller disk. When the thrust coefficient Tc becomes greater than - 0 . 4, the mean velocity through the disk is less than half the free - stream velocity, and the propeller wake is necessarily much larger I

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12 NACA ARR 1'0. 5COl tha.n the propeller diameter . The results of velocity surveys in the tail ::..-egion (figs. 39 and 40) show that the wake in - tensIty and size are a function of T , and the wake location c is a function of ~ . Similar surveys were made with the pro- peller thrust axis M.l ted J but are not presented be caus e the resul ts ' ,,,ere eesentially the same . Consideration of these velocity effects alone would llrooably lead a designer to e x- pect the following results, which are par tially substantiated ye xperim ent: (1) a 10S8 in wing effectivene ss (dCLid~), (2) a loss tn elevator effectiven 0 ss (dc;n!dO ), and (3) loss e in tail load for a given model attitude.

Basi ,c configuration. - As expected , the value of dCLida.

(tail off) with negative thru st appl ied by a single - rotation propeller is 51 porcer-t of the value obtained "lith propellers removed and 43 percent of the ' value ob- tained vii th positive thrust (figs. 16, 17, and 18) . A changG to dual rotation produces an unexplained change in dCLida which raises the foregoing fraction to 60 and 47 percent, respectively. (See figs. 16, 19, and 20 .) These lift-curve slopes are not unreasonable when it is considered that one - third the wing area is behind

the propeller disk, and that at CL = 0.8 (Tc = -0 . 4)

the mean- velocity flow through the propeller is less than half the froe-stream velocity. Also, as the velocity over the wing is reduced, the wing of aspect ratio 5 . 4 is giving the effect of two smaller vlings of aspect ratio 2.

The loss in lift-ctITV C slopo probably could be minimized by Iv,,,ering the wing with respect to the thrust line, by decreasins ,-ling tapar, or by increasing aspect ratio.

Applying negative thrust to tho model with the tail 1'0 -

moved produces a negative increment in em. This increment

groVls vlith C and , at the higher attitudos , the model with L taU off exhibits a n egat:.ve dCm/dCL - A small part of this stability change is due to the ne gativ e- thrust moment, and tl10 remainder must be caused by propeller normal force and the propeller wake effect on wing and fuselage . Evidence given later in this discussion indicates that moments produced by tho normal force are small when the propeller is producing negative thrust .

NACA ARB No . 5COl 13 The complete model operating with negative thrust pos - s e sses moderate stability at 10Yl lift coefficients and very high stability at the higher lift coeffi c ients . The high stebility i s large l y a resLut of a ten d ency towar d stability . with the tail off and the very l ow dCL/d ~ • . With dua l rota - tion the stability is not so great bec a use dCL/d~ is higher th an with single rotation. Contrary to expectation there is no consistent reduction in d Cmt/ d ~ with application of negative thrust) although the los s in elevator effectiveness can be seen by comparing figures 16 , 18 , and 20 .

The cause for the maintenance of dCmt/d~ which is as great as or great er than that ob s erve d with propeller off may be found in the variations of qt/q and ~ (fig . 41) and their subsequent effect on d CIDt/d~ . It can be shown that t he following is true : - dC mt d~ At low angles of attack the stability contributed by the first ter m of the foregoing equation and the high d ~/d~ are nearly sufficient to compensate for the low qt/q . At the 0 ) higher angles of attack (0, >4 the stabilizer moves out of the low- velocity core of the wake (figs. 41 and 42) , and dot/d~ i ncree s es sharply to increase further the stability contributed by the tail. The generally high value of da,t/d~ . can be attributed to the loss in wing d OylDwash behind the

propeller . The sharp increase in d~/d~ between ~ = 4

and a, = 6 is probably a result of the stabilizer moving into the wake up.,ash which is present in the lower half of

high - intensity wakes (reference 2). The same trends of qt

and ~ can be observed in the hinge - moment curves, but , since hinge moments are subject to secondary effects, they do not lend themselves to a direct analysis .

The variation of elevator angle and stick force with speed for the assumed airplane shows adequate stick - fixed and stick - free stability for negative-thrust operation (fig .

42) . For positiv e thrust both the stick - free and stick - fixed -------.J NACA ARR No . 5COl stabiJ.ity are marginal which is not unusual for a highly power- ed airplane of this type .

The minimum t rim speed is about 160 mil es per hour due to the very high stick-fixed stability ( high dCD'lJdCL and 1m.

dCm/d5 ) in the low- speed l'ange . Since this speed is approxi - e matel- the spe 6d f or torpedo launching) such a limitation 'vo l.l ~d be objectionabl e for a torped.o bomber. At first ;it might appear that a compromise in stabilizer incidence .,ould satisfa.ctorily lml er the minimum trim sp eed, but at -Dest only 5 t o 10 miles per hour could be gained because of the e xtreme- l y high stability (dCm/dCL) .

On a fighter - type airplane it would be desirable to change from positive to negative tt ... rust wittlout a change in trim . The chango in e levato r requi.red for trim at a giv8n speed is small in high - spe ed fl ight ivil ere the braking propeller would be us e d.

Tho untrimmed s tic k forces differ bJ 30 or 40 pounds . Proper loca t ion of a trim tab relative t o the prop elle r wake an d olip - 8 tr eam proba bl y could reduce this difference when the tab is so t for trim in pos itive -thrust opera.tion.

Propeller tilt - By til"uing a prop e ll er nose dOi-Ill, the thrllst 1]:;e-{S-iaI :3ed relative to th3 cent er of g~'avity, a.nd the thrust .. U.r.e angle of attack is decr ease d for a given attitu de or lift ccefficient. It l~s been sho~n that for positive thr'l"st the resul tins increments of p:;.-o:?oller-thrust and norILal - force moments ;.,ill increase the stability. I t can be similar- l y reasoned tilat with negative thrust the opposite effect would re realized; that is , the stability would be decreased.

As anticipated} the sale effect of propeller tilt is to shift the rotate the pitc_ing-momcnt curves for either tail on

01' off (figs. 34 and 36). The variations of C2. l q and <Xt

t vlith a, Ilresented in fiCLU'c, 41 show no sicnificant change dU G to propell er tilt . Within the experimental accuracy the change i n moment characteristics can be justified by thv thrust moment ,. hich indicat e s that nermal forces are insIgnificant> Except for possible effects of stalled portions of the b lade, a much reduced normal force should be expected from a propoller d e velopir~ negative thrust, The cause for normal force on a propell er dev e loping positive thrust can be explained NACA ARR No. 5COI b.~ T sketch (a) below vThich shows a propeller-blade . e lement with ito rotational and forw ard velocity vectors.

V(l+a) .

\

./\ 1:/1

eLZ ",/ \~".I / I '" ,r' , "" ;,.' I CLr>eL~ /1 / / . \:" ,If,,' / /' I \ ~.y / / eLp - -\ f' crp eLr / / \ ' , /""" ... I- "7-/. / A ...

//- ,, ' Z ' \' .... r

~

~ ~ . ---~ ;'

21!nr (a) Positive thrust (b) Negative thrust o , Wh0n the prop e ller axis is tilted ap the vector V(l+a) rotates relative to the propeller plan e assuming the position rand Z for blades on the right and left half of the pro- pell e r. The difference betw e en the resulting ry an d, aZ produces unbalanced torque forces between right and left sides of the prop e ller disk. Tho difference between the torqu e for ces is kn oHn as propeller normal forc e . As shmm in sketch (b)) for a prop e ller de livering negative thrust the vector V(l+a) is much small er because a is negativ e . The r e s L ~ting nOl~~l force is very small and in the opposite direc; t ion .

The cha :nge in trim and loss in stick-fixed stability re- sult'i.ng from propeller tilt reduce the minimum trim speed by 5

miles per h 0ur (fi g . 42). This is a step in the right dtrec -

tion b ut still leaves much to be desired. The effects of pro - peller tilt reali~ed with negative thrust) combined with the increased sta b ility for positive thrust, reduce the trim changes between positive- and negative - thrust operation . Use of a trim tab at high speeds might increase the difference in stick forces because of a change in tab effectiveness between positive- and ne gative-thrust operation . Lack of definite criteria sp eci - fying allowable trim changes for a fighter makes final judgment of trim characteristics impossibJe.

Raised horizontal tail. - The change in horizontal tail posi - tion from 1m" to high re sul ted in the following stability increments: (1) an increase at low lift coeffiCients, (2) a

large decrease at C1 = 0 .4 to 0 . 6, and (3) a slight to large

decrease at high lift coefficients depending upon wh ether a single- or dual-rotation propeller was used (figs. 34 and 36).

NACA ARR No . 5COl Tile cause for these incre1llents which produce an S~shape Cm versus C curve can be traced to the variations of qt/q L

and ot (fig . 41). Substitution for each of the variables in

the equation for dCmt/~ (given in discussion of basic con- fil31.J.l"ation characteristics) will sho'w the contribution of each to the stability ",ith a raised horizontal tail . In brief, entry of the tail into the wake is delayed, which changes the

phase relation between qt/q and ot , and which places the

tail in the core of the wake at high lift coefficients when tho '·lake is intense . .

The curves of 0e and stick fo:c-ce versus Vi gi ven in fi8ure 42 (for negative - thrust operation) are not significantly different from those for the basic configuration. Even tho ugh the stabilit y is lower, the minimum trim speed is greater by 5 cr 10 miles per hour because of the very low elevator effective - ness . For positive thrust there is a definite improvement in stability as would be predicted from results of tests on similar models .

Raised horizontal tail and tilted propeller . - Within a fair degree of accuracy, the pitching -mon.ent characterlstics of the model with raised stability and tilted propeller are the pitching - moment characteristics of the basic configuration with the individual e ffe cts of raised horizontal tail and tilte d propell er superimposed ( figs . 34 and 36) . The trim characteristics cannot be obt ained by simple superposition, be caus e raising the horizont.al tail removes any semblance of

linearity in the Sn versus C curves, and tHting the

L p :c opeller rotates the curves bringing a different s e ction of the curves acress the C = 0 ~~is.

m 'The stability of the model as indicated by the variation of 0e and stick force with speed is reduced for negative thrust an d increased for positive thrust . The resulting ' trim changes from positive- to negative - thrust operation are small for any speed greater than 200 miles per hour. The minimum trim spe ed is reduced to 150 miles per hour. Of the four configurations tested this configuration yields the best longitudinal characteristics in steady flightj however) it is apparent t hat the - effects of configuration are relatively small when compared to the basic effects of a propeller prod~cing negative thrust.

NACA ARR No. 5COl 17 Ste~dy fl!g~.0 partia~~ . - To check on possible tail buffeti ng as a result of high negative thrust , a grid of yarns , l ocated in a vertical plane at the tail, VIas observed as the thrust co effi cient was varied. At a Tc of -0.2 the tuft s began oscillating badly, and at a Tc of -0.3 the dis- turban ce increased so that the tufts oscillated through an in- cluded angle of about 45 . For Tc = - 0.4 and -0.5 the flow near the fuse lage r ever sed and the oscillation ,vas so severe th..at it .,ould appear unsafe to operate in this range. Such a test is not necessarily a quantitative measure of buffeting; how ever, it seems likely that full power should not be us ed for braking at low speed .

An obvious solution to the difficulties experienced with full-power braking is to reduce the power output. By properly sel ecting the power and propeller speed it is possible to ob- tain a great variety of Tc versus C relationships (fig. 5 ) .

L The longitudinal - s tabil ity characteristics of the model in the basic configuration are c omp ared in figure 38 for s ev eral power conditions . Only 25 percent power prod u ces half the stability c h..ange (dCmI dC ) between zero pOvler and full power . Zero L pO'\ver appears to be the best operating condition, since it pro- duces gocd braking and yet does not cause large changes in pi tching -.noment characteristics .

With zero power th e longitudinal-stability characteristics for the four con figurations te s ted are nearly the same (figs .

44 and 45). Most of the effect s noted with full pow er are still pr e sent but to a lesser e xtent. Both stick-fix ed and stick··free s tability as sho~m by figur e 46 ar c satisfactory for the pasic configuration, an d, although not pr e sented, were sat- is f actory for all config trr ations tested, The elimination of extreme s tab ility a t low speed and the increas e in elevator effectiveness remove the seri ous limita ti on of minimum trim speed existing with full power. The L\oe required to maintain a given speed upon application of the brake is small, The stick-force increments required to maintain a g iv en speed shml no c on sistent improvement. However , stick-force characteristics are a function of many variables and if L\oe is small the in- c rement s could proba b ly be held within the desired limits by ad ju stm ent of these variables , Accelerated flight .- . One of the flight conditions for which tho use of the propell er as a brake is considered is that of

j

NACA AHH No . SCO l lLmiting the maximum speed in a dive . It is assumed the pj l ot wil l use a predetermined amount of power to hold the desired speed in tho dive and then pul l out \~1thout changing the powe- output of the engine . This maneuver n~y prov e to be t he critical causo of the stick force required of the pilot to produ ce the desired amount of normal accelerat i on" (A steady turn pos sib l y would require more , stick force , but this is not cons1dered to be a normal maneuver with negat i ve thrust. ) The condition sele c ted for analYl3is is a 70 dive at a speed of 310 miles per hour . The thr 'J.st coefficient require d to limit the airplane to this spee d is - 0 . 13 for an assumed drag coefficient of 0 . 025 . The variation of 0e and stick force with normal acceleration for the basic c onfiguratio n of the model is presented in figure 47 . ' In order to give a basis of comparison, the variations of 0e and stick force .. lith normal acceleration in steady turning flight with rated pO ' He):' and posit i ve thrust at the same sp ee d as the dive ' pull - outs are shown in the same figure , The stick--force gra di e nt is only slightly greater for negative - thrust operation (abo ut 18 lb p er g). It wotlid be expected that the use of negative thrust would result in higher stick-force gradients than wHh positive thrust because dCn/dCL is increased and den/doe is decrease d with negative ,thrust . However , dChe/doe de - c reases faster than dCn/do so that ne arly identical stick - e force gradients for positive and negat.ive thrust are the net results . Similar results were , found for the oth er mo del con- figurations . However , the} model ,vi th the raised horizontal tail and inclined thrust axis gave consi d erably higher stick- - forco gradients ( about 28 lb per g) :,"' or both dive pull - outs and steady turning flight .

, The numerical values given for the s tick-force gradients are } igher than are desirable and could be reduced by rede - signing th e horizontal tail , but it is believed that design c hanges for the purpose of ~8ducing the stick - force gradiont with positive thrust ~ill have a simi l ar effect with negative thrust . Since the stick - force gradients are of the same mag ·, ni.tudu for both positive - th..rust and negative - thrust operation , the use of tho propeller a s a dtve brake appears to be satis - factory from the standpoint of sti ck f orce in d ive pull - outs .

NACA ARR No. 5COl Directional Characteristics The directional-stability and directional control charac- t e risti.cs are discussed . und. er the headings Hi . gh speed and. Low

speed . High speed · characteristics ,.ere obtained at av o = -2

which} for positive~thrust operation, corresponds to high-sp eed level flight, anet for negative-thrust operation corresponds to diving flight. Low-speed characteristics were obtained at

au = 9 which, for positive - thrust operation, corresponds to

climbing flight, and for negative-thrust operation corresponds t o decelerating level flight.

High speed. - Inspection of the Cn versus W curves for

the model with tail removed au = _20 (figs . 48 to 52) shows

the medel to be directionally lli1stable with the propeller re - moved and that the application of power Tc = 0 . 03 or -0.019 \ has little e f fec t on the stability. As would be expected from consideration of the velocity effects of the propeller slip - stream, the yawing moment supplied by the vertical tail is increased with positive - thrust and decreased with negative - I [ thr u st operation. - The differences are small, however} and beca u se of small differences in the stability of the model, tail off, the directional stability of the complete model is

I

nearly the same for positive- and negative - thrust opera.tion .

Similar results were observed for the effectiveness of the r u dder (fig . 70(0.)).

LoVl epeed .- The variation of Cn with W for the model

.,ith tail removed au = 9 (figs. 53 to 59 indicates a marked

red u ction in stabilit y for positive - thrust operation Tc = 0 . 33)

particularl y for moderate angles of yaw . Negatlve thrust with

Tc = -0 . 19 has small effect on the stability, but with Tc

= -0 . 38 the stability of t~1e model with tail removed is posi -

tive for the range of angles of yaw between approximately ~100.

For larger angles of yaw the stabilit y becomes negative and approaches the value obtained with th~ . propeller removed. The po sitive stability exhibited for moderate angles of yaw is greater with the single - rotation pr o peller than with the dual- c. ~otation propeller (figs. 68 and 69).

These results may be explained by consideration of the ef - fect of propeller normal force and the effect of the propeller ,- lake on the wing- fuselage combination. Study of the sketch on page 14 will show the effe ' ct of propeller normal force with 20 NACA ARR No . SCOI positive thru st to be d es tabili z ing , and c omp uta tions sho"1" th e effect to be s mall. For negative - thrust operation , the effect of propeller normal force is stabilizing , but of such small magnitude as to be inconsequential. Therefore, it is believe d that the ef fect of the propeller wake on the wing- fuselage com- bination is the principal factor affecting the stability of the mo d el with tail removed . Since the model is unstable with the propeller removed, "it is logical that it should become more un - stable when the fuselage is immersed :n the high - velocity slip- stream associated with positive thrust , and become less unstable when surrounded by the lOvl - veloci ty wake associa ted with nega - tive thrust . The reas on for the S- shape of Cn versus \)J curves fo:i.~ Tc = 0 . 38 ( figs . 68 and 69) is not understood, but it is believed to be primarily an effect caused by the emer- gence of the trailing portion of the fuselage from the propeller walee at large angles of yau .

With the tail on, the directional stability of the mode l for negative-thrust operation is reduced to approximately half that obt~ined with the propeller removed. The stability for

Tc = - 0 . 38 is slightly greater than for Tc = -0.19. This ap -

parently contradictory resu~t is caused by the stability char - acteristics of the model with tail removed . The yawing moment sup~lied by the vertical tail is actually much less for Tc

= - 0 . 38 than for Tc = -0.19, . but because of the positive

stability exhibited by the mode l with tail removed (Tc = -0,38) the resultant stability is slightly greater.

The comparative effectiveness of the rudder (fig . 70(b)) is in the expected direction; that is , the effectiveness is in - creased with positive thrust and reduced with negative thrust .

For negative-thrust operation 'Tc = -0.38 (approximately full

power) the rudder effectiveness is so greatly reduced as to make the operation of the airplane exceedingly unsafe . It is es timated that about 50 of yaw can be produced by use of full rudder with the single-rotation propeller, and about 20 with the dual - ~otation propeller . The reason for this difference in yaw i s not greater rudder effectiveness with the dual - rotation propeller but lower directional stability which may be traced back to the directional characteristics with the tail removed . With Tc limited to - 0.19 (approximately zero power) more than 20 of yaw can be produced by use of full rudder.

IiTACA ARB No. 5COl Rudder hinge-moment characteristics in general follow the same trends as the directional-stability and directional- cont::.~ol oharacteristics, but are not amenable to analysis be- cause of secondary effects.

Some approximate rudder-p eda l- for ce calculations were made neglecting the contribution of the lateral-control system to . directional stability. For the hi~1-speed attitude, au = -2 , and a n indicated airspeed of 300 miles per hour , i t was found that the average gradient of the rudder-pedal force w'as approx- imately the same for positive and negative thrust. Assuming an instantaneous change from positive to negative thrust (Tc = 0.03 to - 0 .19), the change in rudder angle required to hold o

zero sideslip is about 2i r1ght rudder with the si1"'..gle-

rotation p ro p eller , and about lio right rudder with the dual- rotation propeller. The corresponding average change in the tmtrimmed ru dd er -pedal force is about 40 pounds on the right rudder pedal .

For the low-speed altitude (au = 9 ) and an Indicated

airspeed of 160 miles por hour , the change in rudder angle re - qu ired to hol d zero sideslip for the cha1"'~e from positive to

negative thrust (Tc = 0.33 to - 0 .19) is about 23 right

rudder .lith the single-rotation propeller, and about 3 right rudder with the dual-rotation propeller. The corr e sponding changes in rudder-pedal forces are about 120 pounds and 10 pounds on the right rudder pedal, respectively.

The effect of tiltine the thrust axis is to reduce slight- l y both the pedal-force gradient and the change in pedal force f or trim becau s e of the slightl y lower stability exhibited by the model with the tilted thrust axis. No rudder-deflected tests ivere made with the raised horizontal tail be cause of dif- ficulties of deflecting the rudder. However , inspection of the

C and e vorsus \)! c urves for the tail - high configura-

n h r tions (figs. 60 to 69) shows a slight reduction of den/d\)! and dCh)d \)! for anglos of yaw bet.1.,een ±l 6 and a considerable re- duction for greater angles of yaw . This effect may be caused LIS interference between t he horizontal and vertical tail , which in all probability would resttlt in reduce d rudder effectiveness.

From the standpoint of directional stability and con trol , it appears that the bost model configuration is the normal

_J

22 NACA ARR No . 5COl position of tail with the thrust axis tilted down. The use of a dual - rotation propeller ,.ill reduce the trim changes with u se of "f)o"or. For braking operation, the maximum amount of ne8ative thrust which may be used with safety is that corre - sponding to a Tc of about - 0.2 vrhich can be produced with the USB of little if any 1>OI~er (fig. 5 ).

It appears probable that the directional stability and con - trol of a single - engine airplane would be improved for braking operation by the use of twin vertical tails. As one tail en - tered the low~velocity core of the propeller wake, the opposite tail would be emerging into the higher velocity of the free stream . Thus the combined effectiveness of the two tails .;ould tend to remain more nearly uniform throughout the yaw rang e .

Dihedral Effect Tho dihodral effect of the model progressively increases as pOvler is applied to a propel l er producing negative thrust (figs . 71 , 72, and 73) . Convor' sely, the dihedral effect is r educed with ap plication of power to a propeller producing positive thrust. It has been proved that the effect with positive thrust results from the high slipstream velOCity in - creasing the 1if:t on the trailing wtng, and that the effect is a function of wins lift coeffici ent , thrust coefficient , and distance fr om propeller to wing. It can be similarly reasoned that the e ffect with ne gative thrust results from tho low slipstream velocity de creaSing the lift on the trail - ing wing , and that it 1s a function of the same variables . .

Tho increment of dihedral eff e ct can be computed 'Within 20 , percont by assuming that the propeller wake trails in the free-stream direct · ion and that the loss in lift occurs where the wake crosses the .ling .

The large dihedral effect is undeSirable, particularly when couI>led with tho low directional s t ability associated vIi th a braking proI>eller . The lateral - directional corre - spondence, as indicated by the ratio dCnld~ to dC2/d~, var ie s over a wide range when pmwr is changed from full - negative to full - positive thrust (fig . 74) . As a result, if an airplane is designed for proper lat er al-directional corre- spondence in the positive - thrust range (which in itself is a dH'ficult compromise), it ,dll become too s en sitive in roll for high - negative-thrust operation . Since for the divo con- dition both lift and thrust coefficients are small, tho NACA ARB No. SCQl 23 problem is not serious. If full-power braking is used for the torpedo launchIng run, a definite problem exists ; however, otber problems involved in the use of full-power braking at low speed probably will preclude its use. For zero-power braking, the lateral-directional characteristics would not be necessarily id e al but probably would be acceptable .

CONCLUDING REMARKS Tho results of thes e tests indicato that the prop e ller may b0 us ed as a means of spoed control for a single- e ngine tractor-type ai . rplane t o an e xtent which will bc equally or morc eff G ctive than conve ntional spoiler- ty pu dive brak e s, particularly at low spe e ds. The advantag os in f avor of using tho propeller rather than conventional dive brakes, in addi - t ion to gr o at ,~ r cf f o cti veness at low sp eods} are those of simplif i cation of tho aircraft structure and concomitant sav- ing in weight. '1'he disad va ntages ar e the significant changes of tho stability and control characteristics of the airplane produced by the braking propeller. These effects are largely caused by the low-velocity wake of the propeller flowing over wing, fuselage, and tail, rather than any direct forces acting on the propeller.

The results of the tests also show that the undesirable effects of a braking propeller may be minimized by prop er de- s ig n of the airplane. Tilting the propeller thrust axis and locating the tail as r emote from the propeller wake a5 possi- ble help to red u ce the stability and control changes accompany- ing the use of negative thr u st. The use of a dual-rotation pr opeller is of benefit in reducing the changes of rudder angle and rudder-pedal force required for trim with change of power.

By limiting the power output of the engine, thus restrict - ing the amount of negative propeller thrust, the stability and control characteristics may be held within the limits required for safe operation and still produce adequate braking. The possibility of tail buffeting still remains to be investigated q l' antitatively, but visual observation of tufts indicated the absence of serious buffeting for the range of negative thrust corresponding to acceptable operation from the standpoint of stability and control, Ames Aeronautical Laboratory, National Advisory Committee for Aeronau tics, Moffe tt Field, Calif.} Jan. 20, 1945.

24 NACA ARR No . 5COl REFERENCES 1. Hedrick, William S ., and . Dou~la8s, William M.: An Experi - mental Investigation of the Thrust and Torque Produced oy Propellers Used as Aerodynamic Brakes. NACA ARR No, 4H2 6 , 1944.

2. Silverstein, Ate, Katzoff , S., and Bullivant, W. Kenneth : Downwash and Wake Behind Plain and Fiapped Airfoils .

NACA Rep. No. 651 , 1939 .

NACA ARB No. 5C01 25 APPENDIX CONFIGURATION KE Y FOR TEE STABILITY MODEL S b asic config,ll'ation"model in normal fl yi ng condi t ion but vlith out propeller ( i. e . } wing , fuselage J vertical tail } horizontal tail in normal positi on } flaps and gear re- ' tr a cte d) P propell er subscript S denotes single rot a tion s ub script D d enote s dual rotatio n sup erscript d enote s b lade angl e ~ in degr e es at 0.75 radius station H horizontal tail V vertical t ail COEFFICIENTS AND SYMBOLS All coefficients are given i n NACA s tan dard form refe ' rred to the sta Dil itj ax es; and are def in3 d as l'o l101vS : C lift coefficient (L/~ S ) L CD drag c oeff icient (D /~S ) C lateral-f orce coefficient (Y /qS) y C pitching-moment c oefficie nt (M/gSc) ,', m C y a,v ing - moment coefficient (N/ qSD ) n C rolling- moment coefficient (LI/qS b ) C elevator hinge - :n.oment co efficient (He/gSece) h e C rudder h ing e ,-rnom ent co ef f i cient (Rr/gS rc ) h r r 26 NACA ARR No . 5COl where L lift , Ib

D d rag, Ib (also propeller diameter = 2 . 52 ft)

Y cross - wind force, Ib M pitc hing moment , ft - Ib N yawing moment , ft-Ib L ' roll iug moment , ft - Ib He elevator hinge - moment , ft -I b

Hr rudder hinge - moment, ft -l b

T effective thrust, lb

Cl dynamic pressure (~ pye) lb/ sCl ft

S wing area (13.18 SCl ft) c mean aerodynamic chord (1. 627 ft) b vling span (8 . 48 ft) Se el e vator area aft of hir~e line (0.819 sCl ft) c elev a tor cho d aft of hinge line (0.274 ft) e Sr rndder area aft of hinge lin e (0.369 sq ft) c rudder chord aft of hinge line (0.321 ft) r p mass density of air, slugs/cu ft Y airspeed, ft/se c n revolutions per second In presentation and analysis of the results, the follow - ing symbols are used in addition to the coefficients : NACA ARR No. 5COl 27 au uncorrect ed angle of attack of the fuselage reference line, degre e s angle of attack of the fuselage reference line corrected for flow inclination and tm1nel-wall eff e cts, degrees

ut effe ctive angle of attack of the horizon t al tail, de-

grees angl e of y aw of line of symmetry, degr ee s pitching-moment coeffici ent produced by t he tail angle of incidence of propoller thrust axis with respect to fus(;lage reference lin e , degrees angl e of incidence of horizontal tail with r e sp e ct to it f u selag e reference line, degr e es control-surface defl e ction, d eg rees s ub scrip t s e el ev ator I' ru dd er t horizontal tai l CORRECTIONS Th e follcwlng tunnel - vall corrections were applied and are all additive : S :: 6a 57 . 3 X 0ae CLu dC m S 6C -l<- ::

57 .3 -0 - C X x --

m L 3 C u dit *Applied to t ail-on da ta only .

NACA ARR No. SCOl vlh er G

< \ 0.122

8 = 0.135

8 = 0.096 6

C = 70 sq ft

C = uncorrected lift coefficient with tail removeQ

Lu -= -0.031 NACA ARR No . SCOl 29 TABLE 1 .- DIMENSIONS OF STABILITY MODEL Hori z ontal V ertica l Wing tail ta il :_rea , tota l , s<l ft 13 . 181 3 . 007 1 .160 Pp!ln , ft 8 . 479 3 . 667 1. 250 MoaE a s rodynamic ch ord , ft 1 . 627

----- -----

-----------

Aspoct ratio 4. 52 1.345 5 .4 T aper r at i o . 500 . 513 . 59 2 1 . 18 8 Root chord , ft 2. 093 1 . 093 T ip chord, ft 1 .047 .5 61 . 703 NACA 2418 con - NACA 0012- 64 NACA 0012- 64 s t ant to sta - mo d ified to modifi ed to Root section tion 1. 766 feet 10 . 71 per - 8 . 9 pe r cent o utbo ar d of cent thick thic k c enter l in e NACA 0012- 64 NACA 00 12 - 64 mo d ifi ed to modifi ed to Ti p soction NACA 2415 10.71 per- 8 . 9 perc e nt cent thick thick ---------------------------------~--------·------+_------------r - ---------- --; :Jihedral 7 0 wit h ce nte r line . 2 to Incidence with respect to 0 0 l e ft with 2 0 fu s e lage r ef erence line raised hori - ~ontal tail , r ea of movabl e surface aft of 0 . 369 0 . 819 hinge line, B <l ft ------------- 66 . 46 66 . 89 nil~e line , percent c hor d ------------- ,Aero dynamic balance , percent 29 . 9 28.2 of ar ea aft of hinge line ------------- Tai l length, 25 per c ent mean aerod ynamic chord to c ent er 4 .3 15 4 . 320 line hinge , feet --- - --------- Assumod mechanic al ad vantage of c ontro l for com puting

co ntro l forces , f orc e /Ch <l ! ------------- 20

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NACA ARR No. 5001

Fig. 3

(a) Normal tail position and thrust axis.

(b) Raised borizontal tail and thrust axis tilted down 5°.

Figure 3.- Photographs of the stability model mounted in the

Ames 7- by lO-foot tunnel.

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acteri s ti cs of ~tability model in pitch .

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acteristics of s tab il i ty model in pitoh.

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