RM No. L8A23a

I

RESULTS O F PRELlWIlMARY FLIGHT TESTS O F TRE XS-1 AIRPLANE I (8-PERCENT WING) TO A MACH NUMBER OF 1.25 I

By W. C. Williams and ~e E. Beeler

Langley Aeronautical Laboratory Langley Field, Va.

NATIONAL ADVl SORY COMMITTEE FOR AERONAUTICS

WASHINGTON

April 6, 1948 Declassified April 2, 1957

NACA RM No. ~ 8 ~ 2 3 a

NATIONAL ADVISORY COMMITTEE FOR AEXONAUTICS

RESEARCH MEMORANDUM

RESULTS OF PE3LIMINARY FLIGHT TESTS OF THE X S l AIRPLANE

(&PERCENT WING) TO A MACH NUMBER OF 1.25

By W. C. Williams and De E. Beeler

Upon completion of acceptance tests on the XS-1 airplanes by the -- -- -_ - Bell Aircraf t-Corporat tony -one-of- t-hese -a-irplanes (-~~14-which-has-t he thin wing and horizontal tail, 8 percent and 6 percent thick, respectively) was taken over by the Air Forces' Wright Field Flight Test Division for use In an accelerated transonic flight research program. The purpose of these flight tests was to fly at speeds in excess of the speed of sound in as short a test program as possible. No detailed investigations are being made and as large an increase in hch number as compatible vith safety is made in each flight. If necessary, flight will be made at extreme altitudes (50,000 to 60,000 feet). This p ~ o g a n is a cooperative endeavor between the U. S. Air Force and NACA. NACA instrumentation is used in all flights. Data reduction and analysis are performed by NACA personnel. The flying is done by a Wright Field Flight Test Division pilot.

The purpose of this report is to present data from the first flight tests of the XS-1 to speeds beyond a hch number of 1.0. The data pre- sented herein cover a k c h number range from 0.70 to 1.25 and an altitude range from 30,000 feet to 49,000 feet.

AIRPLANE AND INSTRUMENTATION

The XS-1 airplane flown in these tests incorporates an &percent- thi5k wing and &percent-thick tail. Pertinent dimensions of ths airplane ara shown in the three-view layo~t given in figure 1. Flight conditions of the airplane during the tests were as follows:

2 NACA iiM No. ~ 8 A 2 3 s

. . . . . . . . . . . . . . . . . . . . Launchirrg weight, po-x,ia 12,365 La-mching cen t e r4 f+ rav i t y (percent M.A.C.) . . . . . . 22.1 . . . . . . . . . . . . . . . . . . . . . . Lmding weight, pounds 7115 ~ Landir4 c sn t e r4 f -g rav i t y pos i t ion (percent M.A.C. ) . . . . . . . . 25.3

I . . . . . . . . Fuel consumption of eash rocket , pounds per second 7.87 I - I Engine, i'oivr-cylinder RMI-liquid rozket t h ru s t , pounds I p e r c y l i n i e r 1500 . . . . . . . . . . . . . . . . . . . . . . . . .

Mess-urements of airspeed, a l t i t u d e , normal accelera t ion, e levator poeit ion, and t a i l shear leak have been obtained from standard NACA recording i n s t ~ m e n t s i n s t a l l e d i n t he a i rplane. Measurements of a i l e ron posi t ion, s t a b i l i z e r poeit ion, and e levator wheel fo rce were

I , telemetered t o a ground s ta t ion .

I M free--stream Machnumber corrected f o r pos i t ion e r ro r of p i to t - I s t a t i c head

free-stream Mach number uncorrected f o r posi t ion e r r o r of p i to t - s t a t i c head '

-.=

a i rp lane lift .coeff ic ient (measured normal-force component i s assumed t o be equal t o l i f t component ( ~ w / ~ s ) )

dynamic preesure, pounds per foot2

wing area, 130 f e e t 2

horizontal- tai l area, 26, f ea t2

aerodynamic shear load of r i g h t ta i l , pounds -

s tab i l i ze r . inc idence , degrees

e levator posit ion, degrees

angle of a t t ack o f ~ h o r i z o n t a l ta i l , degrees

C ~ t ta i l noncal-force coef f ic ien t (Lp/stq)

NACA RM No. L8A23a

1

TESTS, R-TS, AND DISCUSSION

A ca l ibra t iop of the posit ion e r ro r of the Kollsmari type L i pitot- ! s t a t i c head located 1 chord length ahead of the wing t i p has been mads up t o a corrected Mach number of 1.25. The static-pressure e r rors have

1 been obtained from a survey of t rue s t a t i c pressure within the t e s t a l t i t ude range with the t e s t airplane and using radar t o obtain geometris a l t i tude . The t e s t airplane was flown during the,survey a t speeds where the s t a t i c e r ror was known. The t e s t airplane i s tracked by radar during the t e s t run end the static-pressure e r ror i s determined from a compariaon of the t rue s t a t i c pressure and tha t pressure recorded from the airspeed head of the t e s t airplane. The total-head pressure e r rors have been determined from a theoret ical consideration of the t o t a l head los s behind a deteched bow wave. The cal ibrat ion curve including only the static- pressure errors and the curve including both the s t a t i c and total-head er rors a re noted i n f igure 2. It i s estimated tha t the cal ibrat ion 13

ac-curate-to-a- -M- -of -&Os01--up-t o-a-Mach-number-of-approxi-mat-el-y-142-and-- t o a M of 20.04 above a corrected Mach number of 1.02.

In figure 3 i s shown an envelope of the buffeting region established from lift and h c h number combinations obtained within the buffet region. The boundaries of the envelope have been ident i f ied a s the buffet boundary and l i m i t l i f t . The buffet boundary i s defined by the f i r s t indication of buffet as shown by records of acceleration a.nd wing and t a i l loads. Limit l i f t i s determined during gradual turns where the l i f t ceased t o increase although increasing up-elevator is being applied. The s t ab i l i ze r incidence angle was approximately 2.2O. These data were obtained i n l eve l f l i g h t and i n gradual turns. A n evaluation of the measured t a i l buffeting loads occurring within the envelope shown i n f igure 3 was made. The mixcimum buffeting loads f o r a l t i tudes above 30,000 f e e t were obtained a t l i m i t l i f t from a b c h number of 0.76 t o 0.80 and were of the order of f400 pounds. A t Mach numbers greater than 0.80, buffet loads were l e s s than L250 pounds. As indicated by these l o w buffeting t a i l loads, the buffeting was mild above 30,000 fee t . The p i lo t did not consider the buffeting a serious problem i n negotiating the transonic speed zone.

Figure 4 shows the variat ion of measured quant i t ies with Mach number obtained i n t e s t s made a t approximately 30,000 fee t pressure a l t i t ude f o r a Mach number range from 0.7 t o 0.94. Included on t h i s f igure a re the variations with Mach number of elevator posit ion and force, balancing t a i l - ioad coeff ic ient , and r e l a t ive elevator effectiveness &+/Me. Tests were

made with two s t ab i l i ze r set t ings. The data given i n t h i s figure and subaequent f igures a r ~ f o r essent ia l ly constant l i f t coeff ic ient , With the s t ab i l i ze r set,at an incidence angle of 1.0' the p i lo t did not f l y bsyond a Mach number of 0.876 because it was d i f f i cu l t t o hold steady f l i g h t due t o the elevator forces required fo r trim, the r e l a t ive ly f a r forward posit ion of the wheel with t h i s s t ab i l i ze r set t ing, and because of buffeting expected a t the higher Mach numbers. Data were obtained f o r

4 NACA RM No. ~ 8 ~ 2 3 a

a s t ab i l i ze r incidents of +2.2' up t o a Mach number of 0.934. From these data of elevator required fo r %ria f o r the two s t ab i l i ze r se t t ings e variat ion cf re la t ive elevator effectiveness A u + / A ~ ~ was obtained up

t o a h c h number of 0.876 and i s shown i n t h i s f igure. It shodd be noted tha t the relatitre elevator effectiveness i s reduced by more than 50 percent between a Mach number of 0.70 and 0.87. This reduction i n effectiveness of the elevator w i l l a f fec t the magnitude of the elevator angles required f o r t r i m . It can a iso be seen from the variat iqn of the balancing t a i l load tha t a part of the t r i m change ' is caused by a change i n t h s wing-fuselage moment f o r the Mach number range covered by t h i s figure. These data are i n qual i ta t iye agreement with t e s t s made i n Langley &foot tunnel and wing-flow t e s t s of an XS-1 model.

In f igure 5, the var iat ion of elevator posit ion and force, r igh t a i leron position, and balancing t a i l load with k c h number i s shown f o r a t e s t run made a t 37,000 f e e t pressure a l t i t ude . The maximum value of h c h number reached was approximately 1.00. It should be noted tha t t r i m changes occurred above a k c h number of 0.94 which were i n addition t o those predicted from model t e s t s i n the k c h number range from 0.8 t o 0.94. I n the comparison of the var iat ion of balancing t a i l load and the variat ion of elevator posit ior- with Mach number, several interest ing points are noted. The changes i n elevator posit ion and i n balancing t a i l load a re similar indicating tha t the la rges t e f fec t i s the change i n wing-fuselage moment with Mach number. Also, it should be noted tha t the ,:hange i n t a i l load, indicating change i n wing-fuselage moment b9twean 0.87 and 0.91, corresponds t o a lo change i n elevator position. For the change i n t a i l load occurring near a Mach number of 1.0, which i s appro:.:imately the same magnitude as the e a r l i e r change i n t a i l load,

lo a zhange i n elevator posit ion of approximately 11- was measured. These

2 data indicate a probable fur ther decrease i n elevator effectiveness beyond the change shown i n figure 4. It i s a lso possible tha t some of t h i s elevator deflection i s being used t o of fse t changes i n downwash. The variat ion of r ight a i leron deflection with Mach number shows tha t the airplane i s becoming r ight wing-heavy a s the Mach number increases. The p i lo t reported tha t t h i s wing heaviness was most apparent t o him betwem Mach numbers of 0.90 and 0.92.

The variat ion of elevator posit ion and balancing t a i l load with.Mach number a t 43,000 fee t pressure a l t i t ude up t o a Mach number of approxi- mate ly1 .055 isshowninf igure6 . T h e c u r v e s o n t h i s f i g u r e a r e d i s - continuous because data were selected a t two different values of l i f t coeff ic ient . It can be seen tha t the t a i l load and elevator posit ion follow i n the same manner a s shown i n f igure 5 f o r the same Mach n u d e r range. It should be noted, however, t ha t a t the highest Mach number shown on t h i s figure (1.055), there i s an appreciable reversal i n the direct ion of the elevator motion with l i t t l e or no change i n the t a i l load, indicating possible changes i n the elevator effectiveness or downwash.

NACA RM' No. ~ 8 ~ 2 3 a 5

Figure 7 gives tha var i s t ion of elevator posit ion and forca with Mach number es obtained i n t e s t s mo3e e t a pressure a l t i t ude of 49,000 fee t up t o a Mach number of- approximately 1.25. It should be noted tha t above a bhch number of 1.0, there i s s continuing t r i m change i n the nose-down direction. The m&simum elevator colztrol force required i n f ly ing the X S - 1 i n the transonic speed zone i s shown on t h i s f igure and ocsurs just past a Mach number of 1.0. The force measured was 25 pounds. It should be remembered, however, that these data were obtained a t 49,000 fee t alti- tude. A t lower a l t i tudes , the forces involved i n transonic f l i g h t with the X S l may be greater than the p i lo t can exert . .It should also be pointed out that the XS-1 has a very small elevator. The elevator chord i s 20 percent of the horizontal-taii chord, and the root-mean-square chord

. of the elevator i s only 5.6 inches. With a la rger airplane of similar design the sontrol forces ;nay be unreasonably large.

I n order t o show the e f fec t s of . a l t i tude and' s t ab i l i ze r posit ion on the 1ongitudinal. tr im character is t ics , the var iat ion of elevator position with Mach m b e r from f igures 4, 5 , 6, and 7 i s given i n figure 8. A-l-t.hO*-t

- - - - - e change s i n st aF'IiZeFpTsitiTn3TsmaLl~ 3liEiJ-d-bF

--

pointed out tha t the re la t ive effectiven3ss of the elevator i s low abo-~e a Mach number of 0.8 and it i s expected tha t small changes i n s t ab i l i ze r posit ion may make appreciable difference In the elevator angles f o r t r i m . The data i n t h i s figure show tha t , although the variat ion of elevator m41e with Mach number i s somewhat different f o r each condition shown, . the same general trends a re indicated.

Some d i f f i c u l t i e s have bsen experienced i n recent t e s t s of other airplanes a t transonic speeds with one-dimensional f l u t t e r o r buzz of the ailerons. There has been no evidence t o date of buzz i n the X % l t e s t s . One probable contributing fac tor t o the absence of t h i s osc i l la t ion i n addition t o the t h i n wing section i s the large amount of f r i c t i o n i n the ai leron control system. The f r i c t i o n i n the ailerons i s of the ordar of 20 foot-pounds. The ai lerons a re quite small and even though there i s no aerodynamic balance, the aerodynamic hingg moment of the ai lerons f o r q corresponding t o a Mach nmber of 0.85 and 30,000 fee t , neglecting ef fec ts of Mach number on the hingeeoment coefficient, i s of the order of 7 foot-pounds per degree. Hydraulic dampers a re ins ta l led but have not been used. There a l so has besn no evidence of abrupt cha.ngea i n the f loa t ing tendencies of the ailerons.

CONCLUSIONS

The data obtained i n f l i gh t with the XS-1 airplane with &percent- thick wing up t o and beyond the speed of sound a t an a l t i t ude o f ' , . ' 37,000 fee t and above shuw that most of the t r im and force changes expected i n the transonic range have been experienced. Although-

6 NACA RM No. ~8A23a

conditions are not normal, the airplane can be flown under control through a kch number of 1 at altitudes of 37,000 feet and dove. In detail, the following has been noted:

1. Buffeting has been experienced in level flight but has been mild. The horizontal-tail loads associated with the buffeting have been small.

2. The airplane has experienced longitudinal trim changes in the speed range from 0.8 up to 1.25. The largest control force a~sociated with these trim changes was 25,pounds. The pilot has been able to control the airplane. The relatively small magnitude of the control force may be attributed to the small size of the elevator and the high altitude of the flight.

3. The elevator effectiveness has decreased more than 50 percent in going from a Mach number of 0.7 to 0.87. There is evidence of further reduction in elevator effectiveness above a Mach number of 0.87. This loss in elevator effectiveness has affected the magnitude of the trim changes as noted by the pilot but the actual trim changes for the most part have been caused by changes in the wing-fuselage moment.

4. No aileron buzz or associated phenomenahavebeen experienced. The airplane becomes right wing heavy with increasing Mach number up to a Mach nuniber of 1.10, but can be trimmed with the ailerons.

Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics

Langley Field, Va.

Aeronautical Engineer

Approved : n + I.' * -. , Hartley A:.~oul& Assistant Chief of Research

CMH

NACA RM No. L8A23a

F i ~ u r e 1.- Three view drawing, XS-1 airplane.

NACA RM No. L8A23a

NACA RM No. L8A23a

NACA RM No. L8A23a

NACA RM No. L8A23a

Sub Ject Number

Airplanes - Specific Typss - %11 X S - 1 S tab i l i ty , Longitudinal - S t a t i c S tabi l i ty , Lateral - S t a t i c S tabi l i ty , Longitudinal - Dynamic Controls, Longitudinal Loads, Buffeting and Gust - Tai l

ABSTRACT

Presents r e s u l t s of the U. S. A i r Forces'aascelerated transonic f l i g h t t e s t s of t h s XS-1 No. 1 airplane f o r the Mach number range froru 0.70 t o 1.23 a t a l t i t udes from 30,000 t o 49,000 fee t .

Data a re included on horiz.onta1-tail loads and buffeting, longl- tudinal t r i m changes, elevator gffectiveness and control forces, and l a t e r a l t r i m charac ts r i s t ics .