RESEARCH MEMORANDUM

FLIGBT MEASUREMENTS OF THE LATERAL RESPONSE

CHARACTERISTICS OF THE CONVAIR XF-%?A

DELTA-WING AIRPLANE

By Euclid C. Holleman

? - ? ."""""" UNCLASSIFED

FOR AERONAUTICS WASHINGTON August 5,1955

https://ntrs.nasa.gov/search.jsp?R=19930088805 2018-05-29T06:42:16+00:00Z

3

v

NACA RM ~155~26

NATIONAL

RESEARCH MEMORANDUM

CHARACTERISTICS OF THE CONVAIR XF-W

By Emlid C. Hollman

SUMMARY

As part of the flight research program conducted with the Convair XF-m delta-wing research amlane, rudder pulse maneuvers were obtained at an altitude of about 30,000 feet over a Mach number range of 0.52 to 0.92. Tests were made with and without a wing fence.

- .i

By analyzing these maneuvers the characteristics of the airplane transient, airplane stability derivatives, and frequency-response char- acteristics were measured. The airplane handling qualities were improved by the addition of wing fences. The agreement between experimental and calculated stability derivatives was fair to poor. However by using transfer-function equations from the lateral equations of motion and the experimental stability derivatives, frequency responses were calculated that compared favorably with those determined by Fourier transformation.

Measurements of the aynamic lateral response characterfstics of the airplane were made at an' altitude of about 30,000 feet and over a Mach number range of 0.52 to 0.9 as part of a flight Fnvestigation usFng the XI?-= delta-xfng airplane. Some dynamic lateral response data were also obtained while the effects of w i n g fences on the airplane longitu- dinal characteristics were being investigated. Results of the longitu- dinal stability Fnvestigation with and without wing fences are presented Fn reference 1. The results of simultaneous lateral tests on the air- plane are reported in reference 2; and results of aynamic longitudkml tests are presented in reference 3.

. G

During this phase of the XI?-= test program the dynamic lateral behavior of the airplane was investigated by analyzing the airplane

2 - NACA RM E55E26

. response to abrqpt rudder pulse disturbances. From the recording of each of these maneuvers it was possible to obtain some of the more important stability derivatives and also the frequency-response characteristics of u the aFrplme.

SYMBOLS AND c o m c ~ s

&t transverse acceleration, g units

b w h g span, f%

!e pressure altitude, ft

c2 rolling-mament coefficient

Cn yawing-moment coefficient

CY side-force coefficient

3

-1

=X

=z

= x z

moment of inertia about longit- s tab i l i ty ax is , slug-f t 2

moment 'of i ne r t i a about vertical s t a b i l i t y axis, slug-ft

product of inertia re la t ive to the s t a b i l i t y axis, slug-ft2

2

M

t

v a

B 6r

e

5

cp

9 c

Q

Jr -*

Mach number

time, sec

true velocity, f t /sec

angle of attack, deg

sidesl ip angle, radians or deg

rudder control position, deg

angle between reference axis- and principal axis, posit ive when reference axis is above principal axis a t nose of airplane, deg

damp- r a t i o

r o l l angle, radians

r o l l velocity, rdia,ns/sec

phase angle, deg

y a w angle, radians

4 NACA RM ~ 5 5 ~ 2 6

It yaw velocity, radians/sec

(I) frequency, radians/sec

% undamped natural frequency, radians/sec

Subscripts :

b body axis

The Convair XF"92A airplane is a single-place fighter-type delta- wing airplane powered by a J33-A-29 turbojet engine x i t h afterburner. Physical characteristics of the airplane are presented in table I and a three-view sketch i s presented i n figure 1. For some of the tests a fence w a s located at the 0.607 semispan s ta t ion of the wing. The fence height was equal t o the wing thiclmess at the 0.607 semispan s ta t ion and extended around the w i n g leading edge as shown in figure 2. The airplane inertia in r o l l and yaw about the body axis was obtained from the manufacturer. An inclination of the principal axis of i ne r t i a was estimated t o be lo below the airplane body axis (fig. 1) and the air- plane inertia about the s tab i l i ty axis was calculated for the angle-of- attack range of these tests (fig. 3 ) . Airplane weight and center-of- gravity position were determined f r o m pi lot reports of the amount of fue l remaining a t the conclusion of each maneuver. Average values for these quantities are 13,400 pounds and 27.5 percent of the mean aero- dynamic chord, respectively.

The airplane is controlled by a conventional rudder and by fu l l - span elevons which function as elevators and ailerons. A l l control surfaces are operated by an irreversible hydraulic system w i t h a r t i f i - c i a l f ee l .

INSTRUMEXCATION

Standard NACA recording instrumentation was used to record airspeed, alt i tude, normal acceleration, transverse scceleratfon, yawing velocity, rolling velo.city, angle of attack, angle of sideslip, elevon position, and rudder position. All records were synchronized by a common timer a t

NACA RM -5326 5

? intervals of 0.1 second. An airspeed head, mounted on a boom qprox i - mately 5.4 f e e t ahead of the akrplane nose inlet, m e a s u r e 3 both s ta t ic and t o t a l pressure. Airspeed was calibrated by pacer and radar tracking and is believed t o be accurate t o 20.01 Mach nmber. Control positions were measured by standard control position transmitters and were recorded on a Weston galvanometer which had a flat response t o about 5 cycles per second. Angle of attack and angle of s idesl ip w e r e measured by a vane- type pickup and were also recorded on a Weston galvanometer. The side- s l ip vane pickup d recorder had a flat response t o about 5 cycles per second. Ro l l angular velocity was recorded with a direct recording magnetically damped turnmeter with a natural frequency of X) cycles per second and a damping r a t i o of 0.64. Yaw angular velocity was recorded w i t h the same ty-pe instrument with a natural Frequency of 9.5 cycles per second and a damping r a t i o of 0.67.

d

'TESTS

- The test procedure for this fnvestigation consisted of recording the airplane response t o abrupt rudder pulses. In each instance the afrplane was stabi l ized at the desired test speed and a l t i tude &nd was disturbed by a rapid pulse of the rzldder control. During the disturbance a l l controls except the rudder w e r e ffxed and following the disturbance a l l controls were ffxed u n t i l the airplane returned t o s tabi l ized flight. Figure 4 shows typical h is tor ies of the test maneuver. Tests, w i t h and without a wing fence, w e r e conducted at 30,000 f e e t over a mch nmber range of 0.52 to 0.9.

.r

" H O D S OF W Y S I S

With the present trends in designing high-performance airplanes it has become apparent that motions other than yaw or s idesl ip are important in determfning acceptable dpamic flying qualities . Refer- ence 4 indicated that roll-to-sideslip ratio might be Importmt in p i l o t ra thg of the f lybg qual i t ies of airplanes. In reference 5 the r o l l - to-yaw r a t i o w a s shown t o be useful in determining airplane s t a b i l i t y derivatives. Consequently measurements of the amplitudes of roll, yaw, and sidesl ip have been made f r o m the recorded transients, and have been u t i l i zed in the analysis to give airplane stability derivatives. Ampli- tude ratios and phase relationships of the transient roll ing velocity,

.measured frm recorded t h e histories. However, inasmuch as the recorded time his tor ies a re re la t ive to the airplane body axis, they were converted

done by employing the relation 4 = qb cos a - 'pb sin a. For the

1 yawing velocity, and s idesl ip angle response to rudder pulses have been

'W to s t ab i l i t y ax i s data before p r o c e e w with %e analysis. This was

6 NACA RM ~ 3 5 ~ 2 6

angle-of-attack range of these tes ts it was necessary t o convert only the y a w velocity to the new axis , shce the correct ion to s idesl ip and ro l l ve loc i ty was of the order of 1 percent or less. The order of the correct ion to the yaw velocity is shown in figure 4.

The procedure for determining the amplitude ra t ios a t the airplane natural frequency is graphical in that the free oscil lation record is enclosed by an envelope t o establish the exponential order of the motion. For each maneuver a plot such as figure 5 is obtained from which the

amplitude ra t ios l:l, I f l , and and the time t o dmp t o one-half amplitude a re measured. By careful insepction of the time history, the phase relationships and frequency of the osci l la t ion are determined. The measured amplitude ra t ios and phase angles were converted t o displacement ra t ios by the usual relationships involving “ed natural frequency

and damping angle = & and cp cp = cp - (9” + damping angle). I d lis1 It was shown in reference 5 that the stabil i ty derivatives, CZPJ

C ZP’ c I+’ cnB could be derived from the airplane la teral t ransient

motions. The computing procedure involves the use of an initial approxi- mation for and C , the measured natural frequency, daqping rat io ,

and estimates for the derivatives of lesser importance ( “ 9 9 C yp’ ‘2,)

to calculate the roll-to-yaw amplitude r a t i o and phase angle. The s o h - t ion i s one of i terat ion i n that C2 and C are a l tered until the

calculated amplitude r a t io and phase angle match those measured experi- mentally. When the experimental amplitude r a t io and phase angle are matched, the valuee of C, and \, as w e l l as Cz and C 2 , have been determined.

czB 2P

P 2P

B B P

Reference 6 presents a procedure whereby the a i rplane s tabi l i ty derivatives may be determine& from the airplane frequency-response data by u t i l i z ing a method of l ea s t squares. Sample calculations were made using this method as a check of the resul ts of the previous method.

By means of the Fourier inte@al F(u) =c f (t)e-i&dt the

functions of time were transformed into frequency functions. For this analysis the integral was evaluated by an IBM calculating machine u t i - lizing the method of reference 7. Briefly, the methd of integration f i t s a parabola through the data.ordinates and evaluates the integral by

NACA FM H55E26 - 7

? multiplying the ordinates by a set of coefficients. Summhg these prod- ucts evaluates the integral . -om these calculations the amplitude &

as ra t ios of output t o input and the difference in output to input phase angle.

* phase angle of the complex components w e r e determined and are presented

By using the methods briefly described in the preceding section, the transient-response data have been a n a l y z e d to give airplane stabil i ty derivatives and frequency-response chmacterist ics. The transient- response characterist ics of the airplane at ax^ a l t i tude of 3O,OoO f e e t over a Mach number range of 0.52 t o 0.92 are presented i n figures 6 t o 80 Figure 6 shows the variation with Mach nmber of roll-to-yaw, rol l - to- sideslip, and yaw-to-sideslip amplitude r a t i o a t the natural frequency. The addition of the wFng fence reduced the roll-to-yaw &nd ro l l - to - s ides l ip ra t ios slightly at a Mach nmiber of 0.85. The p i l o t considered this reduction t o be an improvement i n the airplane h a d l i n g qual i t ies a t this t e s t condition. Phase angle relationships were also measured and are shown i n figure 7. Only the amplitude and phase angle of roll t o y a w were used in the present analysis (by the method of ref. 5 ) ; however, the amplitudes and phase angles of r o l l to sideslip and yaw t o sideslip are also presented in figures 6 and 7 t o show the trends. Figure '8 shows the airplane undaruped natural frequency asd damping r a t io fo r t hese t e s t conditions. The measurement of these quantities by the graphical method employed here depends ent i re ly on the airplane response being lightly bmped.

By the met- of reference 5 the more s ignif icant s tabi l i ty derfv-

figure 9. The value of was determined by taking the slope of the

transverse acceleration plotted against sideslip during the airplane's f ree osci l la t ion. The variations of these derivatives with Mach nuuiber a re compared with derivatives calculated by detexdning the l i f t -curve slope of the ver t i ca l tai l (refs. 8 and 9) and by calculating its contri- bution t o the la teral der ivat ives by the method of reference 10. In these calculations, the ver t ica l t a i l area was taken as the area above the fuselage. Wing contributions t o the derivatives were estimated frm the methods of references 8, LL, and E. The wing and t a i l contributions t o the derivatives were sII[rrmed without regard for interference effects. '

The m e a s u r e d sideslip derivatives are compared t o those calculated in figure g(a). Experimental d u e s of are approximately 25 percent

higher than calculated. Thus it appears that the fuselage o r perhaps

B

L

r "ye

8 NACA RM ~ 5 5 ~ 2 6

interference effects contribute a considerable amount to t h i s derivatfve. Experimental Cn shows a dffferent trend than predicted, increasing s l ight ly with Mach number whereas the calculated derivative decreases with Mach number fo r this t e s t range. Trends i n C Z a re similar but

the experimental derivative is approximately one-half the calculated derivative. It appears, then, that the simple,theory used herein i s inadequate in calculating these derivatives. Indicated differences may be the result of influence of the wing wake on the ver t ica l t a i l since these effects were not considered in the calculations. The experimen- tal Cnr (fig. g(b)) is many times larger than the calculated damping i n yaw. A similar discrepancy was noted i n reference 13, particularly a t high angles of attack, and was at t r ibuted to the wing vortex flow creating sidewash over the rear portion of the fueelage. The sidewash lags the airplane oscil lation and Fncreases the tail damping by increasing the angle of attack of the t a i l during the oscil lation. The experimental damping i n r o l l C (fig. g(c)) compares favorably with the calculated value.

B

B

2P

Since the experimental derivatives are functions of the estimated derivatives as w e l l as the measured oscil lation characterist ics of the airplane, calculations were made to indicate the effect of a nominal change i n the calculated derivatives on the experimental derivatives. Results of these calculations are given i n table 11. The maximum effect of changing C by 20 percent appears i n Cnr but this change is only of the order of 5 percent. Altering CZ, changed each of the deriva-

tives but the change was negligible. !Twenty-percent change i n also P al tered each of the derivatives, the maximum change of the order of 5 percent occurring i n C Thus it appears that fairly accurate experi- mental derivatives can be obtained with reasonable estimates for the other derivatives. The estimate of the airplane inertia characterist ics i s also important. For example, the product of iner t ia estimate w i l l Influ- ence C and C, O f course accurate measurements of the motion ampli-

tude ratios and phase angles are necessary. In an attempt t o minimize these errors, faired values for these quantit ies for each Mach number were used in the calculation procedure.

nP

2P

4 B .

Some resul ts of calculating derivatives by the method of reference 6 are a lso included in figure 9. The agreement between the derivatives calculated by the methods of references 5 and 6 is considered good a t the low Mach number but difference6 are apparent a t the higher Mach number, par t icular ly in C and CzP. A measure of the control effectiveness

was also obtained from the method of reference 6 and i s compared t o that measured in t he Ames 40- by 80-foot wind tunnel in figure 10.

4

..

I

J NACA RM ~ 5 5 ~ 2 6 - 9

* By using the Fourier transformation the frequency content of the

transient records has been computed. An example is shown in figure 11. the film record to digital form, error boundaries have been computed as in reference 14 and are also shown in figure 1l. It is evident that at the higher frequencies as the frequency content becakes low, (the expected percentage error becomes h i g h ) the phase angles tend to diverge. Thus accuracy in amplitude assures accuracy in phase angle. This criterion has been used in terminating the fairings of the transfer functions presented.

.r Assuming that the most significant source of error is the reduction of

Shown in figure I 2 is a summary of the frequency-response character- istics of the airplane for four Mach numbers 0.52, 0.63, 0.72, and 0.87 at an altitude of about 30,OOO feet. These data show that the natural frequency and peak amplitude ratio of the airplane increase with increasing Mach number for this Mach nrmiber range.

The results of the transient analysis a n d frequency-response analysis were compared by calculating the frequency-response characteristics of the airplane for the test conditions of figure 12. Transfer-function equations derived f r o m the three lateral equations of motion were used with the experimental stability derivatives and the calculated deriva- tives where experimental derivatives were not available. The inertia characteristics used were fram figure 3. The control effectiveness parameters were obtained from tests of the airplane in the Ames 40- by 80-foot wind tunnel (fig. 10). Results of these calculations at one test Mach ntrmber (0.63) are shown in figure 13. The agreement shown is considered fairly good. Simflar agreement was obtained at the other test Mach numbers.

By analyzing rudder pulse maneuvers with the XI?-= airplane, the characteristics of the airplane transient, airplane stability deriva- tives, and transfer functions were measured. An improvement in the air- plane handling was noted as a result of the addition of the wing fences. Stability derivatives were evaluated experhentally, and were also calcu- lated with fair to poor agreement with experimental data. By us- the experimentally determined stability derivatives, transfer functions were calculated that agreed reasonably well with those cdcuhted by Fourier transformation.

High-speed Flight Station, National Advisory Committee for Aeronautics,

Edwards, C a l i f . , May 18, 1955.

10 NACA RM ~ 5 5 ~ 2 6

I

1. Sisk, Thomas R., Elnd Muhleman, Duane 0.: Longitudinal Stability C h a r - a c t e r i s t i c s in Maneuvering Flight of the Convair XF-W D e l t a - W i n g Airplane Including the Effects of Wing Fences. NACA FM akJ27, 1955.

3. H o l l e m ~ a n , Euclid C., and Triplet t , William C. : Flight Measurements of the Dynamic Longitudinal Stability and Frequency-Response C h a r - a c t e r i s t i c s of the XF-92A Delta-Wing Airplane. NACA lpiI H54J26a, 1955

4. L i d d e l l , Charles J., Jr., Creer, Brent Y., and Van Dyke, Rudolph D., Jr.: A Flight Study of Requirements for Satisfactory Lateral Oscilhtory Characteristics of Fighter Aircraft. NACA FM ~51~16, lggi. .

6. Doregan, James J., Robinson, Samuel W., Jr., and Gates, OrdWeY Bo, Jr-: Determihation of Lateral-Stability Derivative8 and Transfer-Function Coefficients from Frequency-Response Data fo r Lateral Motions. NACA mJ 3083, 1954

7. Schumacher, Lloyd E.: Methoh fo r Analyzing Transient Flight Data t o Obtain Aircraft Frequency Response. (FTD Project No. 49 Rl108). Memo. Rep. No. KRFT-2268, Air Materiel CnmmRnd, Flight Test Div., Air Force, Jan. 17, 190.

8. Fisher, Lewis R.: Approximate Corrections for the Effects of Cam- p ress ib i l i ty on the Subsonic Stability Derivatives of Swept Wings. NACA TN 1854, 1949.

9. %Young, John, and Harper, Charles W.: Theoretical Symmetric Span L o a d i n g a t Subsonic Speeds f o r Wings Having Arbitrary P h n Form. NACA Rep. 921, 1948.

10. Letko, William, and Riley, Donald R. : Effect of an Unswept Wing on the Contribution of Unswept-Tail Configurations to the Law-Speed Stat ic- and Rolling-Stability Derivatives of a Midwing Airplane Model. NACA TJA 2173, 1950.

NACA RM ~55326 ” I1

=? 11. DeYoung, John: Theoretical Antisymmetric Spas Loading fo r Wings of Arbitrary Plan Form at Subsonic Speeds. NACA Fkp. 1056, 1951. (Supersedes NACA !IN 2140. )

t

12. Toll, Thomas A . , and Queijo, M. J.: Approximate Relations and C h a r t s for Low-Speed Stab i l i ty Derivatives of Swept W i n g s . NACA TN 1581, 1948.

13. Johnson, Joseph L., Jr.: --Speed Measurements of Rolling and Yawing Stabi l i ty Derivatives of a 60’ Delta-Wing Model.. W A FM L54G27, 1.954-

14. Cole, Henry A., Jr., Brown, S t u a r t C., and Holl€m3I, &lid C.: merimen- and Predicted Longitudinal Response Characteristics of a Large Flexible 35O Swept-Wing Airplane at an Altitude of 3g,OOO Feet. NACA RM A9H09, lB4.

NACA RM ~ 5 5 ~ 2 6

T m I

PHYSICAL CHARACTERISTICS OF TBE X?-= AlRpLANE

W f n g : Area, sq f t . . . . . . : . . span, f t . . . . . . . . . . Airfoil section . . . . . . . Mean aerodynamic chord, f% . Aspect r a t io . . . . . . . . Root chord, f t . . . . . . . Tip chord . . . . . . . . . . %per r a t i o . . . . . . . . . Sweepback (leading edge), deg Incidence, deg . . . . . . . Dihedral (chord plane), deg .

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

. . . . . . 425 . . . . . . 31.33

. . . . . . 18.09 . . . . . . 2.31 . . . . . . 27.13 . . . . . . 0 . . . . . . 0 . . . . . . fh

65(06)-006.5

. . . . . . U . . . . . . 0

Elevons : Area (total , both, a f t of hinge l ine) , sq ft . . . . . . . . . 76.19

Chord (aft of hinge line, constant except at t i p ) , f t . . . . . 3.05 Span (one elevon), ft . . . . . . . . . . . . . . . . . . . . . 13.35

Movement, deg Elevator: u@ . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5 Down . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

Aileron, t o t a l . . . . . . . . . . . . . . . . . . . . . . . 10 Operation . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic

Vertical tail: Area, sq f t . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.35 Height, above fuselage center line, ft . . . . . . . . . . . . 11.50

R u d d e r : Area, sq f t . . . . . . . . . . . . . . . . . . . . . . . . . . 15.53 Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.22 Travel, deg . . . . . . . . . . . . . . . . . . . . . . . . . . f8.5 Operation. . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic

Fuselage : Length, f% . . . . . . . . . . . . . . . . . . . . . . . . . . 42.80

Power plant: w i n e . . . . . . . . . . . . . . Allison 533-A-29 with afterburner

Rating: S ta t ic thrust at sea level, lb . . . . . . . . . . . . . . 5,- Sta t ic thrust a t sea level with afterburner, l b . . . . . . 7,500

Weight: Gross w e i g h t (560 gal fuel), lb . . . . . . . . . . . . . . . . 15,560 m t y weight, l b . . . . . . . . . . . . . . . . . . . . . . . 11,808

Center-of-gravity locations: Gross weight ( S O g a l f i e l ) , percent M.A.C. . . . . . . . . . 25.5 Bnpty weight, percent M.A.C. . . . . . . . . . . . . . . . . . 29.2

NACA RM ~ 5 5 ~ 2 6

EFFXCT O F VARYING C E R W CALCULATED DERIVATIVES OM THE

Calculated derivatives I m e r

1 - ..

I

"f

3-

c

"

Figure 1. - Three-view drawing of the XF-92A airplane. A l l dimensions inches.

. . .

Foired 69

I b I' b

Station percent chord, c

.. .

16 _I_ NACA RM ~ 5 5 ~ 2 6

'X 3 siug-ft2

J

I

a , deg

Figure 3.- Assumed variation of airplane inertia with angle of attack.

.. . . ..

# c

Right 4

at 0

-4 6

Q, de4l 4

2

Right 2

4 deg 0

2

.. . ..

.

h

Right 2

0 .

2

\

I" "

.-" .I.- " " I

Right I

4, rudadec Q

I

Left 2 I !

s, deg 0- c. I I I

!

I 2 3 4 5 6 7 8 9 ike, t, sec

(b) . M = 0.52; bp 30,000 feet.

~ i g u r e 4.- concluded.

I 0

. .

L

-. . . . . . . . . . . . . . . . . . . . . . . .

-0 3

.02

.01

c

Figure 5.- Verification of the logarithmic order of the airplane oscillatfon. -

- NACA RM 1155326

12. +Basic airplane, clean

II Fence configuration, clean 8

4

0

M

Figure 6.- Amplitude ratio at the natural frequency of r o l l t o yaw, r o l l to sideslip, and y-aw to sideslip f o r the airplane at an a l t i tude of 30,OOO feet (stability ax is ) .

NACA RM ~ 5 5 ~ 2 6

i -180

e+/+, deg -140 -0- Basic airplane, clean a

Fence configuration, clean -I 00

- 360

@+/b, deg -320

-280

Figure 7.- Phase angle relationships at the natural f’reqpency of roll, yaw, and sideslip at a n altitude of 30,OOo feet (stability axis).

22 NACA RM ~155~26

3

2

on, radians/sec

1 -0- Bosic airplane, clean 0 Fence configuration, cleon

0.

M

-1.2 "- Basic airplane, clean Method of analysis

Basic airplane, clean, wing fence] Ref. 5 0 Basic airplane, clean 1 Ref. 6

- - - Calculated - .8

-" ""

- 9

n

- .2- I I

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

"" "_ czp - . I

0

m a u" a

n-

""1 1 ' v

4 .5 R 7 .E3 -9 . 1.0

Figure 9. - Experimental and calculated stability derivatives for the XF-92A airplane.

24 NACA RM E55E26

-4

cnr

"2

0

u Basic airplane, clean Method of' analysis Basic airplane, clean, wing fence Ref. 5

o Basic airplane, clean Ref. 6 "- Calculated

0

E l m "_"__"_ """"".""" """

M

(b ) Yawing derivatives.

Figure 9. - Continued.

J

.2 I

, -i)- Basic airplane, clean Method of analysis Basic airplane, clean, wing fence} Ref. 5

Q Basic airplane, clean 1 Ref. 6 -”

I. Colcuiated . .

% 0 . - ---

“C“

- -””

c

”- -.- /

- .I ~

(c 1 Rolling derlvatives.

Figure 9.- Concluded.

26 II NACA RF4 H55E26

Figure 10.- C o n t r o l effectiveness from the Ames 40- by 80-foot wind- tunnel tests.

NACA RM ~ 5 5 ~ 2 6

cn c EJ W "

E c

.04 -

.OS 1 O Q I IO 1 0

I ' 0 .02 O@cl%- I 1 \ \ \ \ '

.o I \

'\, f l 0 perqent error " " "_

0 -- "" -

Frequency, w , radians/sec

(a) Rudder position.

Figure ll.- The frequency content of a typical run.

28 NACA RM ~ 3 5 ~ 2 6

Frequency, w , radians/sec

(b) Sideslip angle.

Figure 11. - Continued. c

NACA RM ~ 5 5 ~ 2 6

Frequency, to, radians/sec

(c) Yawing velocity.

Figure U.- Continued.

NACA FM ~55.~26

I .6

1.2

I 1 0 0

m8 -lo

\ (' 00 \ \

percent error boundary \ --IO 0

v g-, \

n

\ -. 0 2 . " 4 """"""3"",-

0 . . - - -c- -00 C"

I

Frequency,u, radians/sec

(a) Rolling velocity.

Figure 11.- Concluded.

"

I

Frequency, o, radiansisec

(a) Sideslip angle.

Figure 12.- Frequency reeponse characteristics of the XF-92A airplane at an altitude of 30,000 feet.

- NACA PM H5.5326

M - 0.87 ~. ___- -72 _ _ - .63 "" 52

Frequency, w , radions/sec

(b) Roll ing velocity.

Figure 12.- Continued.

34 - NACA RM ~ 5 5 ~ 2 6

-40

- 80

Q) 0

-0 ” -120 a! > - ai”

8 -200 5

-5 * -160

0 c 0

O

-24 0

-2800 u I 2

Frequency, w, radions/sec

”

.

(a) Yawing velocity about the airplane body axis.

Figure 12.- Concluded.

NACA RM H55E26 - 35

0

-40

-80 mL aP - -120 e a 0 c c3 a

-

-160 E

- 20c

-240

I-

-

I -

-

I-

I-

I- O Frequency, w, radians/sec

(a) Sideslip angle.

Figure 13.- Comparison of the airplane frequency response calculated by Fouzier transformation and f r o m 'afrplane stability derivatives. -

36

20

16

12

a

4

0

-40

-80

I _ NACA RM ~ 5 5 ~ 2 6

-0 -120

cio L

*$-\ 8 Q)

0 E 0

,. -160 -

$ -200 -c a

-240

-280

Frequency, w , rodions/sec

(b) Yawing velocity.

Figure 13. - Continued.

u

i

I

NACA RM ~ 3 5 ~ 2 6 c 37

80 0 Fourier tronsform "_ Calculafed using experimental and

calculated stability derivatives 60

0

40 0 0 @"\

/ 9

/ 9

/ /

\ \

I ',O \

20 ' Q '0,

O- Q * B- 4 -0-0- 0- a-0- 0 0 Q- I

-0 L Frequency, w, radians/sec

(c ) Rolling velocity.

Figure 13. - Concluded.

NACA - Langley Ffeld. Va.