1、NASA TECHNICAL NOTE DYNAMIC STABILITY DERIVATIVES OF A TWIN-JET FIGHTER MODEL FOR ANGLES OF ATTACK FROM -100 TO 1100 . ., Langley Research Center Hampton, Va. 23365 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. JANUARY 1971 Provided by IHSNot for ResaleNo reproduction or networking
2、 permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM 1. Report No. - “ NASA TN D-6091 “ 1 2. Government Accession No. 4. Title and Subtitle DYNAMIC STABILITY DERIVATIVES OF A TWIN-JET FIGHTER MODEL FOR ANGLES OF ATTACX FROM - 10 TO 110 7. Authorts) Sue B. Grafton and Charles E. Libbey “ - .
3、“ 9. Performing Organization Name and Address - “ NASA Langley Research Center Hampton, Va. 23365 . . 12. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington, D.C. 20546 “ 5. Supplementary Notes 1 3. Recipients 5. Report Date Januarv 1971 6. Performing Organiz
4、ation Code 8. Performing Organization Report No. L- 7370 10. Work Unit No. 126-63-11-36 *- 11. Contract or Grant No. 13. Type of Report and Period Covered Technical Note - . . . 14. Sponsoring Agency Code _ . “ .“ - . . 16. Abstract A low-speed investigation was conducted to determine the dynamic st
5、ability deriva- tives in pitch, roll, and yaw over an angle-of-attack range of -loo to 110 for a twin-jet swept-wing fighter model. Several frequencies and amplitudes were investigated to deter- mine the effects of these variables on the stability derivatives. The effect of the vertical and horizont
6、al tail, and horizontal-tail incidence on the derivatives was also evaluated. The results indicate that the model exhibited stable values of damping in pitch over the entire angle-of-attack range, but marked reductions of damping in roll were measured at the stall, and unstable values of damping in
7、yaw were present for the very high angles of attack asso- ciated with flat spins. Either removal of the horizontal or vertical tail or full up deflection of the horizontal tail eliminated the unstable characteristics of the damping-in-yaw derivatives. _ “ - - . . . 17. Key Words (Suggested by Author
8、is) ) Dynamic derivatives Fighter airplane - . 18. Distribution Statement Unclassified - Unlimited 19. Security Classif. (of this report) “ Security Classif. (of this page) Unclassified Unclassified 36 F 21. NO. of Pages 22. Price“ -“ _ “ =. . - For sale by the National Technical Information Service
9、, Springfield, Virginia 22151 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-DYNAMIC STABILITY DERIVATIVES OF A TWIN-JET FIGHTER MODEL FOR ANGLES OF ATTACK FROM -loo TO llOo By Sue B. Grafton and Charles E. Libbey Langley Research Center SUMMARY A l
10、ow-speed investigation was conducted in the Langley full-scale tunnel to deter- mine the dynamic stability derivatives in pitch, roll, and yaw over an angle-of-attack range of -loo to 110 for a twin-jet swept-wing fighter airplane. The study consisted of forced-oscillation tests of a 0.13-scale mode
11、l of the airplane at a Reynolds number of 1.33 x lo6. Several oscillatory frequencies and amplitudes were investigated to deter- mine the effects of these variables on the stability derivatives. The effects of the verti- cal and horizontal tail, and of the horizontal-tail deflection, on the derivati
12、ves were also evaluated. The results of the investigation indicated that the model exhibited stable values of damping in pitch over the entire angle-of-attack range, but marked reductions of damping in roll were measured at the stall, and unstable values of damping in yaw were present for the very h
13、igh angles of attack associated with flat spins. These unstable values of damping in yaw, which constitute propelling moments in a spin, were produced by aero- dynamic interference between the vertical and horizontal tails. Either removal of the horizontal or vertical tail or full up deflection of t
14、he horizontal tail eliminated the pro- pelling moments. INTRODUCTION The National Aeronautics and Space Administration is currently engaged in a research program to develop and validate theoretical methods for prediction of airplane stall and spin characteristics. A major portion of this program inv
15、olves correlation between theoretical calculations and data obtained by free-flight tests using dynamically scaled radio-controlled models of fighter airplanes. Previous theoretical and experimen- tal results obtained for a variable-sweep fighter with a long pointed nose are reported in reference 1.
16、 The present investigation was conducted to measure the dynamic stability deriva- tives of a twin-jet swept-wing fighter model over the angle-of-attack range associated with spinning; these data are intended to serve as aerodynamic input for additional Provided by IHSNot for ResaleNo reproduction or
17、 networking permitted without license from IHS-,-,-. . theoretical studies of the stall and spin characteristics of this particular configuration. The investigation consisted of forced-oscillation tests which were conducted over an angle-of-attack range from -loo to 110 and included the effects of f
18、requency and ampli- tude of the oscillatory motion. Tests were also conducted with the vertical and horizon- tal tails removed. The effect of horizontal-tail deflection on the yawing dynamic deriva- tives is also presented. Pertinent static force tests were conducted to aid in the analysis of the dy
19、namic data. SYMBOLS Physical Concepts The longitudinal and lateral-directional data presented herein are referred to the body-axis system (see fig. 1). All data are referred to a center-of-gravity position of 33-percent mean aerodynamic chord. In order to facilitate international usage of data prese
20、nted, dimensional quantities are presented both in the U.S. Customary Units and in the International System of Units (SI). Equivalent dimensions were determined by using the conversion factors given in reference 2. b wing span, ft (m) - C mean aero9namic chord, ft (m) - Ch mean aerodynamic chord of
21、horizontal tail, ft (m) FA force along X body axis, lb (N) FY force along Y body axis, lb (N) FN force along Z body axis, lb (N) f frequency of oscillation, cycles/sec k reduced-frequency parameter, - 2v wb for lateral-directional parameter or WC - - 2v for longitudinal parameter MX rolling moment,
22、ft-lb (m-N) 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a! pitching moment, ft-lb (m-N) yawing moment, ft-lb (m-N) rolling velocity, rad/sec pitching velocity, rad/sec free-stream dynamic pressure, !?!f lb/ft2 (N/m2) 2 yawing velocity, rad/sec
23、components of resultant velocity VA along X, Y, and Z body axes, respectively, ft/sec (m/sec) free-stream velocity, ft/sec (m/sec) resultant linear velocity, ft/sec (m/sec) body reference axes angle of attack, deg or rad angle of sideslip, deg or rad horizontal-tail deflection, positive when trailin
24、g edge down, deg air density, slugs/ft3 (kg/m3) angle of pitch, deg or rad pitch-angle increment, deg angle of roll, deg or rad roll-angle increment, deg 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Q AQ w angle of yaw, deg or rad yaw-angle incr
25、ement, deg angular velocity, 2nf, rad/sec Dot over symbol indicates derivative with respect to time. Coefficients and Derivatives Results presented herein are given in terms of the coefficients and derivatives defined in the tabulations that follow: . - Normal force aCN CpJ. = - a- 4v2 Eb2 Yawing mo
26、ment 1 “ 2v - 2v cnB = - a- fib aCn 2v Side force cy=- FY qoos cy. =- aCY p bb a- 2v In the present investigation the term “in-phase derivative“ refers to any one of the oscillatory derivatives that is based on the components of forces and moments in phase with the angle of pitch, roll, or yaw produ
27、ced in the oscillatory tests. The term “out-of- phase derivative“ refers to any one of the oscillatory derivatives that is based on the 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-components of forces and moments 90 out of phase with the angle
28、of pitch, roll, or yaw. The oscillatory derivatives of the present investigation were measured in the following combinations : In- phase I I Pitching CmCY - k2Cmq Rolling yi? sin a! - k2Cnb zh . 1 Yawing Out of phase Cmq + Cmk cNq + CNd, + CA a photograph of the model is shown in figure 3, and geome
29、tric characteristics of the full-scale airplane are listed in table I. The longitudinal-control system of the configuration consists of an all-movable horizontal tail which incorporates negative dihedral of 23O to satisfy the lon- gitudinal stability requirements in the normal operational flight ran
30、ge. Lateral control is provided by spoilers as well as ailerons. The ailerons deflect downward while the spoilers deflect upward. The left aileron and right spoiler operate simultaneously as do the right aileron and left spoiler. The directional-control system consists of a conven- tional rudder. Th
31、e maximum control-surface deflections are as follows: Rudder deflection, deg *30 Horizontal-tail deflection (trailing edge), deg 21 up, 9 down Aileron deflection, deg 0 up, 30 down Spoiler deflection, deg 45 up, 0 down All dynamic force tests were made with a single-strut support system and an inter
32、- nal six-component strain-gage balance. The test setups for pitching, rolling, and yawing are illustrated in figure 4 and the equipment and readout system is described in refer- ence 3. The model and the strain-gage balance were mounted on the oscillation sting assembly so that the moment reference
33、 center of the balance was at the center-of-gravity location shown in figure 2 (33 percent F) and was on the axis of rotation for all test condi- tions. Oscillatory motion was imparted to the model by means of a flywheel that was driven by a 3-horsepower (2.2-kilowatt) variable-speed electric motor
34、and a system of pushrods and bellcranks. The amplitude of the oscillatory motion was adjusted by varying the location of the lower pivot point of the vertical connecting rod along the radius of the flywheel. The frequency of the oscillatory motion (limited to about 2 cycles per second) was varied by
35、 changing the speed of the electric motor. A precision sine-cosine potentiometer, which generated voltage signals proportional to the sine and cosine of the flywheel rotation angle, was coupled directly to the flywheel shaft and provided electrical signals proportionai to the angular displacement of
36、 the model. These signals were used in the data readout procedure described in detail in reference 3. TESTS I The forced-oscillation tests in pitch, roll, and yaw were conducted over an angle-of- attack range from -10 to llOo. The tests were conducted in the Langley full-scale tunnel 7 Provided by I
37、HSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-at a speed of 100 feet per second (30.48 m/sec) which corresponds to a Reynolds number of about 1.33 x 106 based on the mean aerodynamic chord of the wing. Tests were made for the complete model and also for the mod
38、el with the horizontal or vertical tails indi- vidually removed, and with the horizontal tails deflected to -21. The tests were con- ducted for amplitudes ranging from *2.47O to *10.5O and at frequencies of 0.5, 0.7, 1.0, and 1.3 and 1.5 cycles per second over the entire angle-of-attack range. RESUL
39、TS AND DISCUSSION The more pertinent results with regard to stall and spin characteristics have been reviewed in previous reports. Lateral-directional characteristics of this airplane at the stall have been reported in reference 4 while aerodynamic factors affecting flat spin ten- dencies are presen
40、ted in reference 5. The results presented herein, which form a com- plete documentation of dynamic stability derivatives at low speed, are therefore presented with a minimum amount of analysis. Static Aerodynamic Characteristics Longitudinal characteristics.- The static longitudinal characteristics
41、of the model as functions of angle of attack are presented in figure 5. Results are shown for horizontal-tail deflection angles of Oo and -21 (corresponding to stick full back), and also for the horizontal-tail-off condition. The normal-force coefficients presented in figure 5(a) indicate that signi
42、ficant flow separation and stall began to appear at an angle of attack of about 12. The variation of the axial-force coefficient shown in figure 5(b) also indicates onset of flow separation at about 12. The results presented in figure 5(c) indicate that the model is statically unstable over a small
43、angle-of-attack range just at the onset of the stall (12 to 18O), but has control effectiveness sufficient to overcome any pitchup or deep-stall trim condition. With the stick full back (6h = -21), the model can be trimmed at an angle of attack well above the stall, and the model is statically stabl
44、e at all elevator positions throughout the spin angle-of-attack region. Above the stall, pitch-control effectiveness decreases to a minimum at about CY = 60 but then increases as angle of attack approaches 90. The results obtained with the horizontal tail off indi- cate that the horizontal tails con
45、tribute a very large diving moment in the spin angle-of- attack region. These results also reveal that the small region of static instability near the stall is also present with the horizontal tail off and is apparently related to wing stall. Lateral-directional characteristics .- The static lateral
46、-directional stability deriva- tives of the model are presented in figure 6 as functions of angle of attack based on values of the coefficients at p = +5O. Results are shown for horizontal-tail deflections of Oo, -21, and (over a limited angle-of-attack range) for the horizontal-tail-off condition.
47、These results indicate that the model is directionally stable up to angles of attack slightly . 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-above the stall (about 21) and that effective dihedral increases about linearly with angle of attack up
48、to a = 12, where flow separation begins, and causes a marked reduction in C until it becomes zero at about the same angle of attack at which the model becomes directionally unstable. This combination of directional instability and zero effective dihedral at angle of attack in the region just beyond
49、the stall is conducive to directional divergence problems which might lead to an inadvertent spin entry. Reference 4 dis- cusses the factors which cause the directional divergence to occur for this airplane con- figuration. The results of reference 4 indicated that loss of directional stability resulted from a combin
