Full report

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3 I i Copy _/<2 .... NASA TM 8X-196 ii ii I I . , TECHNICAL MEMORANDUM SX-196 for the Bureau of Aeronautics, Department of the Navy FREE-SPINNING-TUNNEL INVESTIGATION OF A 1/25-SCALE MODEL OF THE CHANCE VOUGHT F8U-IP AIRPLANE TED NO. NASA AD-3137 By James S. Bowman, 2r., and Frederick M. Healy Langley Research Langley Field, NATIONAL AERONAUTICS AND WASHINGTON Center Va. SPACE ADMINISTRATION - ' 2 i I.;,c) ",W D.- l ilI}i"ftli

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NATION.L AERdN-IY)'¢S _ "_D'fiFAgE :_I_I'RATION TECHNICAL MEMORANDUM SX-196 for the Bureau of Aeronautics, Department of the Navy FREE-SPINNING-TUNNEL INVESTIGATION OF A 1/29-SCALE MODEL OF THE CHANCE VOUGHT F8U-1P AIRPLANE* TED NO. NASA AD-3137 By James S. Bowman, Jr., and Frederick M. Healy SL%_t_RY An investigation has been made in the Langley 20-foot free-spinning tunnel on a 1/29-scale dynamic model to determine the spin and recovery characteristics of the Chance Vought F8U-1P airplane. Results indicated that the F8U-1P airplane would have spin-recovery characteristics similar to the XF8U-1 design, a model of which was tested and the results of the tests reported in NACA Research Memorandum SL_6L31b. The results indicate that some modification in the design, or some special technique for recovery, is required in order to insure satisfactory recovery from fully developed erect spins. The recommended recovery technique for the F8U-1P will be full rudder reversal and movement of ailerons full with the spin (stick right in a right spin) with full deflection of the wing leadingedge flap. Inverted spins will be difficult obtained should be readily terminated the yawing rotation and neutralization controls. to obtain and any inverted spin by full rudder reversal to oppose of the longitudinal and lateral In an emergency, the same size parachute recommended for the XF8U-1 airplane will be adequate for termination of the spin: a stable parachute 17.7 feet in diameter (projected) with a drag coefficient of 1.1g (based on projected diameter) and a towline length of 56.5 feet. *Title, Unclassified. L-714

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2 wo Dw. . INTRODUCTION ,, io _,,u Q At the request of the Bureau of Aeronautics, Department of the Navy, an investigation has been made in the Langley 20-foot free-spinnlng tunnel of the spin and spln-recovery characteristics of a 1/2-scale model of the Chance Vought F8U-1P airplane. Figure 1 is a three-vlew drawing of the model as tested. The F8U-1P model is similar to the XF8U-1 model previously tested in the spin tunnel (ref. l) except that the lower fuselage forebody cross section has been modified to accommodate the camera installation. Figure 2 illustrates the nature of the modification. Spln-tunnel tests on a 1/29-scale Eodel of the XF8U-1 airplane indicated that both a flat rapidly rotating spin and a steeper slower oscillatory spin were possible (ref. 1). Subsequently, however, static force tests at high angles of attack and tests of a 1/9-scale dynamic radio-controlled model (ref. 2) of a similar design indicated that the flat fast spin was a result of low Reynolds number. The present investigation was undertaken because references 3 and 4 indicated that the F8U-1P cross-sectlonal shape was such that a propelling pro-spln yawing moment would prevail on the fuselage nose at angles of attack of 70 ° or higher for both model and airplane Reynolds numbers and that a flat rapidly rotating spin as well as a steeper slower spin would likely be possible on the corresponding airplane. The erect and inverted spin and recovery characteristics of the model were determined with the model loaded to simulate the basic flight design gross weight (center of gravity at 23..9 percent _) of the airplane. Erect-spln tests were also made with the center of gravity at 51.9 percent . The influence of the gyroscopic mo_ents of the rotating engine components on erect spins and recoveries wa investigated. Brief tests were also made with a spln-recovery parachute housing simulated on the model. A spin-recovery tall parachute to effect satisfactory spin recovery in an emergency was also investlga_ed. SYMBOLS b wing span, ft Cy side-force coefficient, i c mean aerodynamic chord, ft Fy side force, lb Fy 1 0(v)2s

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.... . w v . ..... Ix, Iy,Iz moments of inertia about X-, Y-, and Z-body axes, respectively, slug-ft 2 Ix - Iy inertia yawing-moment parameter mb 2 Iy - I Z inertia rolling-moment parameter mb 2 Iz - IX inertia pitching-moment parameter mb 2 Z fuselage depth at wing root leading edge, ft mass of airplane, slugs R Reynolds number (based on fuselage depth at wing root V'Z leading edge), -- V S wing area, sq ft V full-scale true rate of descent, ft/sec V' absolute velocity, ft/sec x/_ ratio of distance of center of gravity rearward of leading edge of mean aerodynamic ratio of distance between reference line to mean chord to mean aerodynamic chord center of gravity and fuselage aerodynamic chord (positive when center of gravity is below line) angle between fuselage reference line and vertical (approximately equal to absolute value of angle of attack at plane of symmetry), absolute angle of attack, sideslip angle, deg deg deg m relative density of airplane, pS--_

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4 .......... 6_FID_IAL - . ...... kinematic viscosity of air, st_md_rd condition, ft2/sec P air density, slugs/cu ft angle between span axis and horizontal, deg full-scale angular velocity about spin axis, rps MODEL AND TESTING TECHNIQLrES The 1/25-scale model of the Chance Vought F8U-LP airplane was constructed at the Langley Research Center of the National Aeronautics and Space Administration. The dimensional characteristics of the airplane are presented in table I. The mass characteristics for the loadings of the airplane and for the loadings tested on the model are presented in table II. The model was ballasted airplane at an altitude of 50,000 to obtain dynamic similarity to the feet (p = 0.000889 slug/cu ft). A remote-control mechanism was install_d in the model to actuate the controls for the recovery attempts. Sufficient torque was applied to the controls for the recovery attempts t( reverse them fully and rapidly. Controls were set with an accuracy of ±l °. The angular momentum of the rotating components of the full-scale engine was simulated by rotating a flywheel with a small battery-powered motor. The flywheel was located in the mod_l so that the axis of rotation was parallel to the longitudinal axis cf the airplane. Tests were made with and without the flywheel The following normal maximum dicular to the control hinge lines) Rudder, deg: Right ............................. Left .............................. Horizontal tail (trailing edge), deg: rotating. control d_flections (measured perpenwere used during the test program: 6 6 Up ............................... 50 Down .............................. Ailerons, deg: iO Up ........................... 15 Down ........................... 15

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..... GONF.ID AL........ General descriptions of model testing techniques, methods of interpreting test results, and correlation between model and airplane results are presented in reference 3. Model spin-recovery information as presented in chart 1 includes the following notation: For recovery attempts in which a model strikes the safety net while it is still in a spin, the recovery is recorded as greater than the number of turns from the time the controls are moved to the time the model strikes the net, as out control movement (rudder held_rlth recorded as "no spin." >3. When a model recovers withthe spin), the results are RESULTS AND DISCUSSION Erect Spins Basic flight design gross weight.- The results of tests with the model loaded to simulate the basic flight design gross weight (loading 1 in table II) are presented in chart 1. to the right and to the left indicated Inasmuch as the results for spins no significant effects of model asymmetry, the data are arbitrarily presented in terms of right spins. Recoveries from erect spins of the model were generally attempted by simultaneous reversal of the rudder movement of the ailerons to full with to full against the spin, and the spin (stick to the right in a right spin). Selection of this procedure as the normal control recovery technique was based on the results of XF8U-1 model tests reported in reference 1 and on the effectiveness of control techniques in terminating spins of airplanes having various conditions of mass distribution as discussed in detail in reference 3. The spins in which the ailerons were either neutral or against the spin during the developed phase of the spin exhibited two spinning conditions - a flat rapidly rotating spin and a steeper more oscillatory spin; in some instances, also, an additional condition in which the model would not remain in a developed spin was also obtainable. When the ailerons were maintained full with the launching rotation, the model would not spin. The criterion spin configuration indicated that model recoveries could range from satisfactory to unsatisfactory. Based on the model results obtained, it is considered that satisfactory airplane recoveries may not always be obtained by the normal control recovery technique (rapid rudder reversal to full against the spin and movement of the ailerons to full with the spin). During airplane recovery attempts the stick should be maintained full back, inasmuch as the model results indicate that faster rates of rotation recovery more difficult. When r6covery r at forward stick positions make appears i_nlnent, the stick should

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6 - ........ -COl_ID_fI_I, - ....: ..: be moved forward to prevent entry into a sec)ndary spin in the opposite direction. To insure satisfactory plane, modifications to the design XF8U-1 in reference 1 are necessary. recovery haracteristics for the airsimilar to those recommended for the A special control technique (discussed later) should provide satisfactory recovery from at least the steep-type spins. Other conditions.- Erect spin and recovery characteristics of the model were investigated with the center of gravity moved from 25.9 percent _ (loading I in table II) to 31.9 percent S (loading 4 in table II). Tests were made with the angular momentum of the rotating components of the engine at idle rpm simulated by a flywheel mounted in the model. Clockwise and counterclockwise rotations of the flywheel were investigated in both right and left erect spins. Brief tests were made with a spin-recovery parachute housing simulated on the model (fig. 1). No significant variations in the spin and recovery characteristics of the model as reported for the basic flight design gross weight were observed for any of these conditions. Special recovery technique.- Model spin tests conducted on the XF8U-I (ref. l) indicated that the extension of canard surfaces on the nose of the airplane would provide satisfactory reccveries from either the flat or the steeper type of spins when used in ccnJunction with the optimum control technique. However, during the air;lane spin demonstration (ref. 5) in which only steep-type spins were obtained, the contractor elected to utilize full (landing) wing leading-edge-flap deflection in conjunction with the optimum control manipulation. Satisfactory recovery characteristics have been indicated from steep-type spins of the XF8U-1 by using this control technique. It should be pointed out, however, that, based on the data presented in referezce l, this control technique would not be sufficient to provide type spin. However, in the absence satisfactory recoveries from the flatof the ecommended canard modification, full (landing) deflection of the leading-edge flaps (as used on the XF8U-1) in conjunction with the previously specifie recovery technique (simultaneous rudder reversal to full against the spin and aileron movement to full with the spin) is recommended of the airplane. in attemsting recovery from erect spins Inverted Spins The results of inverted spin tests wit the model loaded to simulate the basic flight design gross weight (loadirg 1 in table II) indicated that the model would not enter a developed inverted spin for any condition of control-surface settings investigated. Based on the model tests, it appears that the F8U-1P airplane would b_: difficult to spin inverted. In the event that an inverted spin is encotu:tered with the airplane, recovery should be satisfactory by the met_>d recommended for the XF8U-1

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• m _ !:!7_ .! B .......•• in reference 1 - that is, full reversal ........ 7 ° v ..... .... of the rudder to oppose the yawing rotation and neutralization of the longitudinal and lateral controls• Spin-Recovery Parachute Tests Brief tests utilizing the spin-recovery parachute that was found to be satisfactory for the XF8U-1 of reference 1 were conducted. Results of these tests indicate that the same parachute would provide satisfactory spin recovery for the F8U-1P airplane during emergencies in spin demonstrations. These tests were conducted for the basic flight design gross weight (loading 1 in table II). The towline was attached to the bottom of the extreme rearward point of the fuselage. The rudder was maintained full with the spin during the recovery attempts. The parachute was a 17.7-foot-diameter (projected) stable parachute with a drag coefficient of 1.14 (based on projected 37.5 feet long and the towline length area). The shroud lines were was 36.5 feet. Another size stable tail parachute giving equivalent drag could also be used for satisfactory recovery. Effects of Reynolds Number and Tunnel Testing Technique Reynolds number and tunnel testing technique may have considerable effect on the spin-recovery results of some contemporary fighter design models tested in the spin tunnel. Experience has indicated as pointed out in references 3 and 4, that the part of the fuselage forward of the wing (hereinafter referred to as the nose) can introduce autorotative or antirotative moments depending on the cross-sectlonal shape of the nose and on the Reynolds number. The technique used in testing the models in the spin tunnel (ref. 3) involves launching the models in a flat attitude with rotation. This technique provides favorable conditions flat spin as well as a steeper spin. for the model to find a possible The corresponding airplane, on the other hand, may be capable of simulating such an entry only as a result of a violent maneuver, a pitch-up, or a directional divergence. However, as a general case_ the airplane enters the spin from a low-angle-of-attack, no-rotation condition from which it is difficult to increase the rotational rate to the fast flat spinning condition. In order to evaluate better the current model spin results, static force tests were made on a 1/9-scale model of the F8U-1P nose (in the presence of the rest of the fuselage). The force tests were made in the Langley 300-MPH 7- by lO-foot tunnel for a range of Reynolds numbers to represent both the model and airplane. The results of these tests for the spinning angle-of-attack range of the model are presented in figure 3 as the variation of side-force coefficient with sideslip for a range of

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Reynolds numbers. As indicated in figure 3, the F8U-LP nose design is expected to have little or no Reynolds number effect. The side force on the nose through the spinning angle-of-attack range indicates that an antirotative or damping moment is produced on the nose of both the spin model and the airplane for angles of attack up to 60 ° . However, from 70 ° and up, an autorotative or propelling moment is indicated as possible on both the spin model and the airl_lane. It appears, therefore, from the foregoing discussion that the two types of spins obtainable on the model are also possible on the F8U-1P airplane, but it is expected that the steeper type would be most likely to be obtained unless some violent maneuver leading to pitch-up and directional divergence should It is recommended that intentional occur on the F8U-1P configuration. spins be avoided with the F8U-1P airplane inasmuch as there is no assurance that even the special technique of utilizing full wing leading-edge-flap deflection will be effective if the spin should develop to the flat phase. In the event of an inadvertent spin or a violent maneuver likely t(, induce a spin, recovery should be initiated immediately to minimize the possibility of entering the flatter type of spin. SUMMARY OF RESULTS From a free-spinning tunnel investigation of a i/2_-scale model of the Chance Vought F8U-IP airplane, applicable to the spin and recovery 30,000 feet: the foll_wing results are considered charact_ristics of the airplane at i. Two types of erect spins are possible with the ailerons neutral or against the spin: a flat rapidly rotatiiLg spin, and a steeper more oscillatory spin. Satisfactory recovery may sometimes not be possible from either type of spin even by the normal control recovery manipulation (simultaneous rudder reversal of ailerons to full with the spin) to full _gainst the spin and movement and some airplane modification appears necessary to insure satisfactory recovery f:'om all developed spins that are obtainable on this design. As an altemlative, the special technique employed for the XF8U-I airplane (full defl,_ction of wing leading-edge flap in conjunction with the normal control technique) may be effective for insuring recovery. It is recommended that the spin not be allowed to develop fully on this airplane and that :'ecovery control technique be utilized as soon as a spin is indicated. 2. Center-of-gravity movement gyroscopic moments of the rotating tion of the spin-recovery parachute from 23.!) percent _ to 31.9 percent _, components of the engine, or installahousing on the lower rear fuselage

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had little significant influence on the erect spin and recovery characteristics of the airplane. 5. Inverted spins are difficult to obtain. The recommended recovery procedure if an inverted spin is encountered is full reversal of the rudder to oppose the yawing rotation and neutralization of the other controls. 4. For satisfactory emergency spin recovery during demonstration flights the 17.7-foot-diameter (projected) stable tail parachute with a drag coefficient of 1.14 (based on projected area), 37.5-foot shroud lines, and a 36.5-foot towline previously utilized for the XF8U-1 airplane is adequate. Langley Research Center, National Aeronautics and Space Administration, Langley Field, Va., August 28, 1959. REFERENCES 1. Klinar, Walter J., Lee, Henry A., and Wilkes, L. Faye: Free-Spinning- Tunnel Investigation of a 1/25-Scale Model of the Chance Vought XF8U-1 Airplane - TED NO. NACA DE 392. NACA RM SL56LSlb, Bur. Aero., 1956. 2. Libbey, Charles E., and Burk, Sanger M., Jr.: A Technique Utilizing Free-Flying Radio-Controlled Models To Study the Incipient- and Developed-Spin Characteristics of Airplanes. NASA MEMO 2-6-59L, 1959. 3. Neihouse, Anshal I., Klinar, Walter J., and Scher, Stanley H.: Status of Spin Research for Recent Airplane Designs. NACA RM L57F12, 1957. 4. Polhamus, Edward C.: Effect of Flow Incidence and Reynolds Number on Low-Speed Aerodynamic Characteristics of Several Noncircular Cylinders With Applications to Directional Stability and Spinning. NACA TN 4176, 1958. 5. Alley, J. R.: Model XF8U-I/F8U-I Airplane Part II Spin Demonstration Plus Additional Investigation and Demonstration of Spins With the Aileron Control Held Full Against the Spin - XF8U-1 Bureau No. 138900. Rep. No. 10638 (Contract Nos. N0a(s) 54-605, 53-1075), Chance Vought Aircraft, Inc., Jan. 15, 1957.

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lO TABLE I DIMENSIONAL CHARACTERISTICS OF THE CHANCE VOUGHT FSU-1P AIRPLANE Overall length, ft ....................... Wing: 94.25 Span, ft ........................... 35.67 Area (including chord-extension), sq ft ........... _89.55 Root chord, in ........................ 202.00 Tip chord (not including chord-extenslon), in ......... 49.95 Tip chord (including chord-extenslon), in ........... 95.93 Mean aerodynsmic chord, in .................. 141.40 Distance from leading edge of _ rearward of leading edge of root chord, in ....................... 92.20 Aspect ratio (area including chord-extension) ........ 5.50 Taper ratio (not including chord-extenslon) ......... 0.29 Taper ratio (including chord-extenslon) ........... 0.28 Dihedral, deg ........................ -5 Incidence, deg ........................ -i Sweepback at quarter-chord, deg Airfoil section: Root .......................... Tip .......................... Ailerons: ............. 42 NACA 65A006 NACA 65A005 Total area (rearward of hinge llne), sq ft .......... 85.12 Span of one aileron, percent b/2 .............. 58.55 Horizontal tail: Span, ft ........................... 18.09 Area, sq ft ......................... 93 .45 Sweepback at quarter-chord, deg ............... 45 Root chord, in ........................ 108.05 Tip chord, in ......................... 15.96 Aspect ratio ......................... 5.55 Dihedral, deg ........................ 9.42 Airfoil section: Root ..................... Tip ..................... Vertical tail: Modified NACA 65A006 Modified NACA 65A004 Height, ft .......................... 12.08 Total area (including dorsal), sq ft ............. 115.95 Rudder area (rearward of hinge line), sq _t ......... 12.56 Sweepback at quarter-chord, deg ............... 45 Root chord (51 in. above fuselage reference llne), in ..... 155.00 Tip chord, in ......................... 41.00 Aspect ratio ......................... 1.26 Airfoil section: Root (69 in. above fuselage reference llne) ............. ModlfiedNACA 65A005.3 Tip .................... Modified NACA 65A004

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...... _OI'IF_I_kL'" m x II _ , 7 i i _,o,g ,s I ,,-4 : ."........ 11 - - .... o x x x x x k x x × _, o ,s ,s ,s mcx H ,ll I I I ,-4

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12 -"......... tD_"I'DE]TIAL" .'. ..... " " - - - - . 4" • CHART 1.- CT SpTn D RIC0_Y CT_L,TICS OF T MODwr- [RaOOVO attired by e_t_amoo ov_eal of x_dd to ull aLnat the spin movmnt o£ alle_m to ftLll with the spin unless othomw_o diestod ep) (FeO@wl_°y attwtod omp _ dffllO_Jd lpk M p_Olmo to, Ft_4IF 1 with a_ane : Attitudes tpM dMtl Loadin_ MU-1P oat R_ht Bu£e altitude Des_m_d 30,000 ft b I Z91 0.35 33h o. 78 ?U7D Kor_mo1_1 tail 2/3 up :, j d [ 81 12u 52 26UI &tle_o_e 72 28D! full a83 o.b.l_ 313 o.z? (St_ek 5, JT_ , i m) SFIU d I b k 1 83_ 6U 62 12U z z . _ so sPz a_W oonditiona sJible. bvea'y osalllato Ip qn • (see table II) iI ht design Uos8 wo_ht o_mto-of-av| poslttm: 23,9 pememot e b • 76 9_._ ,_ o.,_ r b k I 79 3U _ IU *ller_s ] [ ] _ I &n 5 zTD 1 i ] full I ] I I b I 8o 7U 1 5PDII . 5 ,. (q) ( l] V n • £t4 z_OOVO_y, model entered s IIp_n in the oppoel_e dlz'*otlmo. Model Values Omo- (tp•) (m) dTl_ee oe_d_tim poeeLble, 4- tim values. Tm-u For °Model emt4wed • 8ho_t d£ve £ollove4 b7 • spin oppos£te dlraotlmo. fF_dal mote&'od a Klide. V_0d t0 oo_J_olilpond£ full-seal, U _ _ u] -eo0_017 D nordO_D wLa4 lReoover'y attired b7 s_ts +c_s r_srsal of rudd_ to 2/3 s4P l_t the spin and movemMt of tlor o 2/3 with t spin. h£f%_ r'l, oo'Ty, model t_nod in _ oppo£t• d£z_,otlc_. _L)Iodol l_llod InVea._d and oDtoa'od a gl£do. JF_dol _ooove_od Ln an Lnvewted CL_o. IC)d41 eazte_od • dlvo. 1AFt4 eo_vezT, model ant•re4 mo alle *'oll. n_lol entered an lnvez.ted dive.

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_CONFZDENT.13J-'" : :'" _ :.... 15 . i, • w _ - .... V Figure i.- Three-view drawing of the 1/29-scale model of the Chance Vought F8U-1P airplane. Center-of-gravlty position indicated is for the basic flight design gross weight.

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14 -. .w_ _ oe eoe oe --,-A I F81J-IP £orwaz'd A_els_e PSu-1P Section A- A XF8U-I Figure 2.- Typical cross sections of the F8U-LP and XF8U-I forward fuselages.

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C_NF_TL_J_ ......... 19 .2 .I a':80°-_ Cy 0 _ --.i I J I 2 0 5 I0 B,deg R 0 403,000 [] 570,000 < 806,000 Autorotative "_ ct'=60 o Anti rotative , I i J 15 20 Figure 3.- Effect of Reynolds number and angle of attack on the autorotative tendencies of the F8U-1P fuselage (forward of wing-fuselage intersection).

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NATIONAL TECHNICALMEMORANDL_8X-196 for the Bureau of Aeronautics, Department of the Navy FREE-SPINNING-TUNNELINVESTIGATIONOFA 1/29-SCALEMODEL OF THECHANCEVOUGHTF8U-IP AIRPLANE* TEDNO. NASAAD-3157 By James S. Bowman3 Jr., and Frederick M. Healy ABSTRACT Results of an investigation of a dynamic model in the Langley 20-foot free-spinning tunnel are presented. Erect and inverted developed spin and recovery characteristics were investigated. The size of stable tail parachute required for spin recovery in an emergencywas determined. INDEXHEADINGS Airplanes - Specific Types Spinning Mass and Gyroscopic Problems Parachutes Piloting Techniques *Title, Uncl__, .!__# 1.7.i.2 1.8.3 J',_ 1.8.6 i. i0 7.7

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