1、c RESEARCH MEMORANDUM DYNAMIC LONGITUDINAL STABILITY CHARACTERSSTICS OF A SWEPI“w7NG FIGHTER-TYPE AlRPLAJXE AT,MACH I I I 1 - NUMBERS BETWEEN 0.36 AND 1.45 Yx=k a 4“ By Chester H, Wolowicz High-Speed Flight Station I al Li I I t I ! Edwards, Calif. Cl t t . L 2 t I rz c- 2, m FOR AERONAUTICS WASHING
2、TON April 1, 1957 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EEACA RM m6K03 c SWEPT-WING FIm-TYPE AESIATKE AT MACE Ey Chester E. Wolawfcz As part of the flight research program conducted by the Natianal Advisory Committee for Aeronautics on a sw
3、ept-wing fighter-type -lane not equipped with an automatic pitch damper, pulse mmeuvers were per- formed at altitudes from TD,OOO to 40,000 feet over a Mach nmiber range from 0.36 to 1.45 to determine the longitudinal stability character- istics and derivatives for an original-ang and an extended wi
4、ng-tip configuration. The longitudinal dynamic behavior of the airplane during sinniLated combat maneuversat altitudes of 30,oOO to 40,000 feet was not considered satisfactory, especially at supersonic speeds, because of insufficient pitch aRmping. The addition of the wing-tip extensions caused a sl
5、ight favorable shift in the aerodynamic center of the airplane. The static =gin of the extended wing-tip configuration is of the order of l2-percent mean aerodynamic chord in the subsonic region aad 29-percent mean aer0aynarmtc chord at Mach numbers above 1.2. Wind-tunnel data for the two wing confi
6、gurations investigated showed good agreement with transonic flight results for the Lift-curve slope and the static stabilfty derlvative %; poor aweement was evident in the supersonic region. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM W
7、HO3 INTRODUCTIOW - The static and dynamic longitudinal stability characteristics ELnd derivatives, as determined fram flight pulse data, for two wing con- figurations of a 450 swept-wing fighter-type airplane capable of rright well into the superscnic region axe presented in this paper. Stabilizer p
8、ulse data employed were obtained Par an original-wing configuration and also for a configuration with a 1-foot extermion of the wing tip. All data were obtained within the 10,000- and 40,000-foot levels over the Mach nrmiber ramge from 0.36 to 1.45 at the NACA High-Speed Flfght Station at Edwards, C
9、alif. The results of the flight data analysis are compared with available wind-tunnel data which have been corrected for the momentum effects of the intake air of the jet engine. This paper constitutes one part of a general flight investigation of the stability, performance, and aerdymmic load chara
10、cteristics of the dr-plane. Results of 60me other investigations have been reported in references 1 to 4. /sec weight of airplane, Ib angle of attack of airplane, angle between reference body axis and the relative wind, per radian Fn equations, per deg in figures rate of change of angle of attack wi
11、th time, radians/sec inboard slat position, percent of fully open position outboard slat position, percent of“-filly open position ratio of actual damping to critical damping mass density of air, sluga/cu A- ft The test airplane is a fighter-type with a 45O swept wing and a low horizontal tail. It i
12、s powered by a single turbojet, engine equipped with an afterburner. A three-view drawhg of the airplane wlth the orig- inal vertical tail is shown in figure 1. Figure 1 also shows a dotted outline of the wing employed in the exbended-wing configuration. A photo- graph of the airplane is sham in fig
13、ure 2. The wing-tip extensions were added to increase the static margin and improve the stability for the external wing-mounted fuel-tank configuration. The airplane was not equipped wlth an automatic pitch damper. The data for the original-wing and extended wing-tip configuretiom were obtained with
14、 several different vertical tails mounted on the air- p-e at vasioue tines during the tests (ref. 4). The effects of the changes in the vertical tails on the lon;itudinal stability character- istics are considered negligible. The airplane is equipped wlth automatic leading-edge slats installed as fi
15、ve interconnected segments. At 40,000 feet, the slats were open at bhch numbers below 0.84 for steady flight; the slats started to apen n response to air loads at angles of attack of bo, 5O, 7O, and 8O, at Mach numbers of 0.84, 0.9, 1.03, and 1.08, respectively. At 20,000 feet, the slats were open a
16、t Mach numbers below 0.72 for steady fUght; the slats started to open at angles of attack of 4 and 6 at Mach numbers of 0.72 and 0.86, respectively. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The physical characteristics of the two configuration
17、s we presented in table I. The estimated mation with airplane weight of the moment of Inertia relative to the pitch axis (fig. 3) is based on the manufac- turers estimate for design weight and empty weight conditions (ref. 5) . Standard NACA instruments were used. to record airspeed, altitude, pitch
18、ing velocity and acceleration, normal acceleration, angle of attack, control-surface poeitions, and leadling-edge slat positions. The angle of attack, airspeed, and altitude were sensed on the nose boom. All records were synchronized st 0.1-second intervals by a common timing circuit. The pitch turn
19、meter used to measure the pitching velocity and acceleration is cansidered accurate to within M.5 percent of range. The turnmeter mounting direction error is 0.50 or less. The indicated normal acceleraneter readLngs were corrected to the center of gravlty. The accelerometer is considered accurate to
20、 within m.5 percent of range. The vane-type pickup for measurSng the angle of attack was EBBS balanced and had dynamically flat response characteristics over the frequency rmge of the airplane. Although the pfckup is statically accurate to fO.lo, the indicated angle of attack has been corrected only
21、 for pitching velocity to the center of gravi.ty of the airplane. The ranges, aynamic chaxacteristics, and scales of recorded data for the angle-of-attack, velocity, and acceleration instruments are: Scale of I Undaurped I recorded data natural (per in. I frequencies, 1 am ping ratio deflection) cps
22、 I 10.0 to 10.55 o .70 8 o.gg to 1.075 0.65 14 1.38 to 2.16 0.65 7 to 8 4.48 to 5.93 0.38 at 40,000 ft 0.43 at 30,OOO ft 0.48 at 20,000 ft 0.3 at 10,OOO ft 19 0.33 at p,oOO ft Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 7 NACA RM H56HO3 t Contr
23、ol-surface and leading-edge slat positions were measured by standard control-position transmLtters. The control-surface position transmitters were linked directly to the control surfaces and are con- sidered accurate to within M .lo. The nose-boom installation for measur- the airspeed was calibrated
24、 by NACA radar phototheodolite method. The Mach nunibers presented are considered accurate to a.02. The test procedure for this investigation consisted of recording the airplane response to abwt stabilizer pulees performed Hith the other controls fixed. In all instances the pilot attempted to mainta
25、in constant hch number and altitude and to prevent movement of the control surfaces during the transient portion of the maneuver. Figures 4(a) and 4(b) present typical time histories. L The stabilizer pulse maneuvers were generally performed at lg f 0.lg % condi-t;ions; however, for the original win
26、g configuratioa at Mach numbers above M EZ 1.05 the maneuvers were performed at various load factors and altitudes from 40,000 to 35,OOO feet. Pulse weuvers at Mach num- v bers greater .t;han 1.35 were performed following a pull-out *om a dive. The following table lists the altitudes and carrespondi
27、ng Mach number ranges for which data were obtained for each conffration: Configuration Mach number range Altitude, ft original wing 0.36 to 0.93 10,000 0.77 to 1.49 40,000 40,000 0.53 to 1.03 Extended wing 0.79 to 1.26 - ANALYSIS A preliminary study of the data showed no aigniflcmt nonlinear influen
28、ces, hence linearized, emall disturbance, short-period forms of the longitudiqal equations of motion of the airplane constituted the basis of the analysis. The time-vector method of analysis (refs. 4, 6, 7, and 8) was employed to determine the derivatives. Because of the lack of reliability of the d
29、etermined values of ( % + k), this quantity is not presenteu. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MCA RM H56E03 7 . “ . The magnitudes of C and (cm, + C%), as determined by the time- % vector method of analysis, were spot-checked by usi
30、ng the following equa- tions and were found to be in agreement. The original-wing area was employed in analyzing all flight data. To convert the derivatives of the extended-wing configmatian to the actual wlng mea and wing-chord basis, the Cr, derivative should be dtipUed by 0.98, C., by 0.9, and (%
31、 4- qy) by 1.01. In faFring the test points to obtain a constant altitude, lg curve, consideration was given to the influence of altitude and load factor on the test points when the test points were obtained from maneuvers at other than the desired altitude and load factor conditions. A summary of t
32、he figurea presenting the results of this investi- gation is: . “ . - 1 Configuration lkfluence of wing- tip extensions 5 5 - - Figures l0ngituW derivatives Comparison with rating criteria 9 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The varia
33、tion of trfm a with Mach number sham in figure 5 for three distinct altitudes has been included not only to show trim a but also to aid in estimating the probability, during the pulse maneuvers, of the automatic opening of the slats when use is made of the information previously presented in the sec
34、tion describing the airplane. Original Wing, On the basis of available data the period curve (fig. 6) shows a smooth and normal large decrease in the Mach number reglon between 0.85 and 0.95, followed by a more gradual decrease to the highest Mach num- ber. The damp- r therefore, fairing the T1/2 po
35、ints has not been attewted. Insufficient eta in this region precluded the possibility of defining a reliable curve. . ma, The magnitudes of and C and the variation of these deriv- ma atives with Mach nurriber (fig. 7) show generally good agreement with wind- tunnel datal (ref. 9) which were correcte
36、d Tor the momentum effects of the intake sir of the Jet engine. It should be pointed out that in the Mach number region between 0.6 and 0.g there is appreciable-scatter of Ch points, considerably above the experimental. scatter, which may be in accordance with the rapid variations with Mach number s
37、hown in Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM H56H03 9 I . L I references 10 and U. It was not possible to verify the presence of the rapid variations in % from a study of available Kind-tunnel data because of the lack of wind-tunne
38、l test points wfthin this region. Extended wing The results of the analysis of the data for the eXtenaea-wing air- plane (figs. 8 and Lo) show the same generel behavior of the individual quantities plotted as functions of Mach nmiber as was discussed for the original-wing configuration; consequently
39、, detailed consideration of the variation of the quantities with Mach nuniber fs omitted. The decrease in period wbich occurred wfth decrease in altitude (fig. 8) is primarFly due to the corresportding increase in dynamic pres- sure, overshadowing the effect of decreasing %, which would tend to incr
40、ease the period. If the aerodynamic derivatives of equatfon 3 were invariant with altitude, the damping ratio 5 could be expected to increase as the altitude is decreased. The increase Fn f with decrease in altitude at subsonic speeds, as shown in figure 8, fs considerably less thsn would be obtaine
41、d by a in a.ir density alone. This condition is attrib- utable to the decrease Fn the etude of the wing derivative (“p + k) wLth decreasing altitude. Pilot opinion indicated that the airplane, which did not have a pitit wer, was unsatisfactory insofar as the longitudinal mc behavior was concerned du
42、ring simulated carbat at altitudes varying from 40,000 to 30,000 feet. At supersonic Mach nmibers, the atrplane had initial rapid and abrlxpt response to control input follawed by prolonged, rapid short-period oscillations. At low subsonic Mach numbers, the air- plane had a slow initial response fol
43、lared by prolonged slow oscillatione which required concentration to eliminate. The most acceptable, but still unsatisfactory, characteristica were noticed in the vicinity of M = 0.8. me results of the analysis have been plotted on a qualitative rating chart (fig. 9) obtained from reference 12; pilo
44、ts opinion showed good qualitative agreement with the criteria of figure 9. Caution should be ased in nor . do wlnd-tunnel data and incomplete flight data indicate any significant influence of slats on CL,. - The c.0, curves (fig. LO) shbw distinct altitude effects prfmarlly in the region of the tra
45、nsonic aerodpmnic-center shirt. A study of the unpubUshed Langley the increase in tends to increase the damping ratio. The reason for the apparent negative decrease in (% + c poor agreement was evLdent in the supersonic re#on. High-Speed Plight Station, NationaX.-Advisory Corninittee for Aeronautics
46、, =wards, Calif., ly 23, 1956. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 REFEREMCES NACA RM E36H03 1. MCA High-speed Flight Station: Flight Ekperience With Two High- Speed Airplanes Having Violent Lateral-Longitudinal Coupling in Aileron Rol
47、ls. NACA RM H55Al3, 1955. 2. Finch, Thomas W., Peele, James R., and Day, Richard E. : FUght Invest- igation of the Effect of Vertical-Tail Size on the Rolling Behavior of a Swept-Wing Airplane Eaving Lateral-Longitudinal Coupling. MACA RM H5528a, 1956. 3. Drake, Hubert M., Finch, Thomas W., and Peel
48、e, James R. : Flight Measurements of Directional Stability to a Mach Number of 1.48 for an Airplane Tested With Three Different Vertical Tail Canffguratione. NACA RM 3526, 1935. “ 4. Wolawicz, Chester E.: Time-Vector Determined Lateral Derivatives of a Swept-Wing Fighter-Type Airplane With. Wee Diff
49、erent Vertical Tails at Mach Numbers Between 0.70 and 1.48. NACA RM 5620, 1956. 5. Weight Control Section, N.A.A.: Airplape Inertia Calculations for an Air Superiority Fighter Airplane - Day, (Monoplane), AF Model F-100A, (NU We1 NA-l32), Contract Bo. AF33( 600) -6543. Rep. No. NA-52-185, North American Aviation, Inc., Oct. 1953. 6. Breuh
