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Large scale wind tunnel tests to improve longitudinal stability of supersonic transports with variable swept wings
Some problems relating to longitudinal stability in power-on flight are considered. A derivation is included which shows that, under certain conditions, the rate of change of the pitching moment coefficient with lift coefficient as obtained in wind tunnel tests simulating constant power operation is directly proportional to one of the indices of stability commonly associated with flight analysis, (the slope of the curve relating the elevator angle for trim and lift coefficient). The necessity of analyzing power-on wind tunnel data for trim conditions is emphasized, and a method is provided for converting data obtained from constant thrust tests to simulated constant throttle flight conditions.
Three matrice methods are developed and presented for determining the longitudinal-stability derivatives from transient flight data. In these methods the expressions for some of the stability derivatives are in the form generally used in stability calculations. The first method requires the combination of four measurements in time-history form, two of which must be incremental elevator deflection and incremental tail load and the other two measurements can be chosen from a possible three, namely incremental load factor, pitching velocity, and angle of attack. The method demonstrates the use of the tail load to separate the pitching-moment derivatives and to determine the downwash derivative. (author)
The subsonic longitudinal stability and control derivatives of the F-18 High Angle of Attack Research Vehicle (HARV) are extracted from dynamic flight data using a maximum likelihood parameter identification technique. The technique uses the linearized aircraft equations of motion in their continuous/discrete form and accounts for state and measurement noise as well as thrust-vectoring effects. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics over the aircraft, particularly at high angles of attack. Thrust vectoring was implemented using electrohydraulically-actuated nozzle postexit vanes and a specialized research flight control system. During maneuvers, a control system feature provided independent aerodynamic control surface inputs and independent thrust-vectoring vane inputs, thereby eliminating correlations between the aircraft states and controls. Substantial variations in control excitation and dynamic response were exhibited for maneuvers conducted at different angles of attack. Opposing vane interactions caused most thrust-vectoring inputs to experience some exhaust plume interference and thus reduced effectiveness. The estimated stability and control derivatives are plotted, and a discussion relates them to predicted values and maneuver quality.
A matrix method is presented for determining the longitudinal-stability coefficients and frequency response of an aircraft from arbitrary maneuvers. The method is devised so that it can be applied to time-history measurements of combinations of such simple quantities as angle of attack, pitching velocity, load factor, elevator angle, and hinge moment to obtain the over-all coefficients. Although the method has been devised primarily for the evaluation of stability coefficients which are of primary interest in most aircraft loads and stability studies, it can be used also, with a simple additional computation, to determine the frequency-response characteristics. The entire procedure can be applied or extended to other problems which can be expressed by linear differential equations.
An investigation of the static and dynamic longitudinal stability characteristics of 1/3.7 scale rocket-powered model of the Bell MX-776A has been made for a Mach number range from 0.8 to 1.6. Two models were tested with all control surfaces at 0 degree deflection and centers of gravity located 1/4 and 1/2 body diameters, respectively, ahead of the equivalent design location. Both models were stable about the trim conditions but did not trim at 0 degree angle of attack because of slight constructional asymmetries. The results indicated that the variation of lift and pitching moment was not linear with angle of attack. Both lift-curve slope and pitching-moment-curve slope were of the smallest magnitude near 0 degree angle of attack. In general, an increase in angle of attack was accompanied by a rearward movement of the aerodynamic center as the rear wing moved out of the downwash from the forward surfaces. This characteristic was more pronounced in the transonic region. The dynamic stability in the form of total damping factor varied with normal-force coefficient but was greatest for both models at a Mach number of approximately 1.25. The damping factor was greater at the lower trim normal-force coefficients except at a Mach number of 1.0. At that speed the damping factor was of about the same magnitude for both models. The drag coefficient increased with trim normal-force coefficient and was largest in the transonic region.
This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.
Low speed wind tunnel test on effects of horizontal tail geometry on longitudinal stability of orbiter shuttle
An evaluation of the longitudinal stability and control characteristics of a mixed-flow remote-lift (MFRL) STOVL aircraft in the powered-lift portion of the flight envelope is presented. A stabilization and command augmentation system was implemented on the MFRL aircraft to meet the requirements for satisfactory flying qualities. The pitch portion of this control system uses a state-rate feedback implicit model following controller to achieve the desired flying qualities and to suppress the effects of external variations and disturbances in the aircrafts characteristics over the low speed envelope.
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The effect of restraint in pitching introduced by an automatic pilot upon the longitudinal stability of an airplane has been studied. Customary simplifying assumptions have been made in setting down the equations of motion, and the results of computations based on the simplified equations are presented to show the effect of an automatic pilot installed in an airplane of known dimensions and characteristics. The equations developed have been applied by making calculations for a Clark biplane and a Fairchild 22 monoplane.
Tests were conducted at Mach numbers of 3.96 and 4.65 in the Langley Unitary Plan wind tunnel to determine the static longitudinal stability characteristics of a fin-stabilized rocket-vehicle configuration which had a rearward facing step located upstream of the fins. Two fin sizes and planforms, a delta and a clipped delta, were tested. The angle of attack was varied from 6 deg to -6 deg and the Reynolds number based on model 6 length was about 10 x 10. The configuration with the larger fins (clipped delta) had a center of pressure slightly rearward of and an initial normal-force-curve slope slightly higher than that of the configuration with the smaller fins (delta) as would be expected. Calculations of the stability parameters gave a slightly lower initial slope of the normal-force curve than measured data, probably because of boundary-layer separation ahead of the step. The calculated center of pressure agreed well with the measured data. Measured and calculated increments in the initial slope of the normal-force curve and in the center of pressure, due to changing fins, were in excellent agreement indicating that separated flow downstream of the step did not influence flow over the fins. This result was consistent with data from schlieren photographs.
A fixed-base simulator study was conducted to determine the minimum acceptable level of longitudinal stability for a representative turbofan STOL (short take-off and landing) transport airplane during the landing approach. Real-time digital simulation techniques were used. The computer was programed with equations of motion for six degrees of freedom, and the aerodynamic inputs were based on measured wind-tunnel data. The primary piloting task was an instrument approach to a breakout at a 60-m (200-ft) ceiling.
In the present report, aerodynamic relations under the effect of the weight moments are investigated, and an example given of the computation of the stability for a practical case. Later, the effects of the masses of the controls on the dynamic longitudinal stability will be considered.
Effects of wing-crank, leading-edge chord extensions & horizontal-tail height on longitudinal stability of swept-wing models
A brief investigation was made of the longitudinal-stability characteristics of a YF-84A airplane (Army Serial No. 45-79488). The airplane developed a pitching-up tendency at approximately 0.80 Mach number which necessitated large push forces and down-elevator deflections for further increases in speed. In steady turns at 35,000 feet with the center of gravity at 28.3 percent mean aerodynamic chord for normal accelerations up to the maximum test value, the control-force gradients were excessive at Mach numbers over 0.78. Airplane buffeting did not present a serious problem in accelerated or unaccelerated flight at 15,000 and 35,000 feet up to the maximum test Mach number of 0.84. It is believed that excessive control force would be the limiting factor in attaining speeds in excess of 0.84 Mach number, especially at altitudes below 35,000 feet.
The possibility of obtaining useful estimates of the static longitudinal stability of aircraft flying at high supersonic Mach numbers at angles of attack between 0 and +/-180 deg is explored. Existing theories, empirical formulas, and graphical procedures are employed to estimate the normal-force and pitching-moment characteristics of an example airplane configuration consisting of an ogive-cylinder body, trapezoidal wing, and cruciform trapezoidal tail. Existing wind-tunnel data for this configuration at a Mach number of 6.86 provide an evaluation of the estimates up to an angle of attack of 35 deg. Evaluation at higher angles of attack is afforded by data obtained from wind-tunnel tests made with the same configuration at angles of attack between 30 and 150 deg at five Mach numbers between 2.5 and 3.55. Over the ranges of Mach numbers and angles of attack investigated, predictions of normal force and center-of-pressure locations for the configuration considered agree well with those obtained experimentally, particularly at the higher Mach numbers.
Results of an investigation of the static longitudinal stability and control characteristics of an aspect-ratio-3.1, unswept wing configuration equipped with an aspect-ratio-4, unswept horizontal tail are presented without analysis for the Mach number range from 0.70 to 2.22. The hinge line of the all-movable horizontal tail was in the extended wing chord plane, 1.66 wing mean aerodynamic chords behind the reference center of moments. The ratio of the area of the exposed horizontal-tail panels to the total area of the wing was 13.3 percent and the ratio of the total areas was 19.9 percent. Data are presented at angles of attack ranging"from -6 deg to +18 deg for the horizontal tail set at angles ranging from +5 deg to -20 deg and for the tail removed.