A summary of the longitudinal and lateral stability and control characteristics obtained from rocket-model tests of a swept-wing fighter-type airplane at Mach numbers from 0.5 to 1.9

A f l i g h t investigation has been conducted by means of rocket models of a swept-wing fighter-type airplane t o determine drag coefficients, longitudinal.and l a t e r a l s t a b i l i t y derivatives, effects of aeroelasticity on r o l l i n g effectiveness, and the effect of the engine j e t exhaust on the trim characteristics over the Mach number range from 0.5 t o 1.9.The jet-engine simulator caused a decrease i n trim angle of attack of approximately 1.270 and a decrease i n t r i m -l i f t coefficient of 0.07.A positive increment i n pressure coefficient was caused by the j e t on the side and bottom of the fuselage.As the distance downstream of the j e t e x i t increased, the increment on the bottom of the fuselage increased, whereas the increments on the side decreased t o a negative peak.The drag r i s e begins a t a Mach number of 0.90.The minimmi-drag coefficient (including base and internal drag) has a value of 0.02 a t a Mach number of 0.87, an increase t o 0.070 a t a Mach number of 1.1 and then a gradual increase t o a value of 0.074 a t a Mach number of 1.83.There was a reduction i n s t a t i c longitudinal s t a b i l i t y at the higher l i f t coefficients a t subsonic speeds.Dynamic longitudinal s t a b i l i t y was indicated throughout the speed range.The horizontal tail was an effective control throughout the speed range.The dihedral effect was adequate.The r o l l -ing was nearly constant through the speed range and agreed with some theoretical values.The aeroelastic losses i n rolling effectiveness varied from about 6 percent a t 35,000 feet t o about 27 percent a t sea l e v e l a t a Mach rimer of 0.5 and from about 20 percent a t 35,000 f e e t t o about 84 percent a t sea l e v e l NACA RM ~5 6 ~1 9As a result of the current interest in airplanes that fly at supersonic speeds, the Pilotless Aircraft Research Division of the Langley Aeronautical Laboratory has conducted an investigation to determine some of the aerodynamic characteristics of a twin-engine, swept-wing, fightertype airplane by utilization of the rocket-propelled-model technique.The primary purposes of this investigation were to determine drag coefficients, longitudinal and lateral stability derivatives, effects of aeroelasticity on the rolling effectiveness, and the effect of the engine jet exhaust on the trim characteristics, since the engine exits are located below and well forward of the all-movable horizontal stabilizer and tail.This paper summarizes the flight-test results obtained from the six models used to determine the desired aerodynamic information over the Mach number range from 0.5 to 1.9.SYMBOLS A cross-sectional area, sq ft A, jet exit area, sq in.a total damping factor longitudinal-accelerometer reading wg normal-accelerometer reading "t/g transverse-accelerometer reading b wing span, ft E mean aerodynamic chord, ft c chord-force coefficient, positive in a reasward direction, a2 W 3 -g Sw 9 C~ drag coefficient, CN sin a + CC cos a NACA R b I ~5 6 ~3 .9 base-drag coefficient, --PO) C ~, b base area q%~ ' D, i internal-drag coefficient C ~, m i n minirmnn-drag coefficient hinge-moment coefficient, Ringe moment qStEt C~ l i f t coefficient, CN cos a, -Cc s i n a / c l i f t coefficient f o r minimum drag coefficient ~)cD,min \ %I pitching-moment coefficient about center of gravity pitching-moment coefficient about center of gravity a t zero o angle of a t t a c k and horizontal-tail deflection cm = acm/i($), per raiiian 9 cmk = >C /a(%), ZV,: per radian Cm + C pitch-darrrping derivative 9 a C~ normal-force coefficient, positive toward top of model from an W 1 model center l i n e , ---€5 S , q incremental change i n pressure coefficient due t o power-on, Cp,power-on -Cp,power-off pressure coefficient, (PI -PO) coefficient of r o l l i n g moment due t o r o l l i n g velocity, acz e r radian dig)' c o e f f i c i e n t of r o l l i n g moment due t o yawing v e l o c i t y , ac 1 -, p e r radian ac 2 coefficient of r o l l i n g moment due t o s i d e s l i p , -, per r a d i a n &P c o e f f i c i e n t of yawing moment due t o r o l l i n g velocity, ' cn -, p e r r a d i a n c o e f f i c i e n t of yawing moment due t o yawing velocity, ac c o e f f i c i e n t of yawing moment due t o s i d e s l i p , 2, p e r r a d i a n aP c o e f f i c i e n t of yawing moment due t o sideslipping velocity, side-force c o e f f i c i e n t c o e f f i c i e n t of s i d e force due t o s i d e s l i p , 3, p e r radian dB t h r u s t , l b a c c e l e r a t i o n due t o gravity, 32.2 f t / s e c 2 moment of i n e r t i a about body r o l l a x i s , slug-ft2 moment of i n e r t i a about body p i t c h axis, s l u g -f t 2 mor-ent of i n e r t i a about body yaw axis, slug-ft2 product of i n e r t i a , slug-ft2 length, f t Mach number