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X-15 aircraft static longitudinal, directional, and lateral stability and control at hypersonic speeds
This report contains the flight-test results of the longitudinal-stability and -control phase of a general flying qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). The tests were conducted at indicated airspeeds up to 530 miles per hour (0.76 Mach number) at low altitude and up to 350 miles per hour (0.82) Mach number) at high altitude. These tests showed that the flying qualities of the airplane were in accordance with the requirements of the Army Air Forces Stability and Control Specification except for excessive elevator control forces in maneuvering flight and the inadequacy of the longitudinal trimming control at low airspeeds.
In the area of control, several problems can arise during the evolution of the space station. These incude: 1) the use of multiple or articulated flexible bodies; 2) the need for distributed cntrol for maneuvering and maintaining altitude; 3) hierarchical control to automate and manage control systems; 4) structual control (from the standpoint of appendage stamping; isolation, and possible figure control); 5) cntrol position and orientation for component modules during construction (an evolutionary requirement); 6) control during docking and undocking operations; and 7) the normal requirements for stablity and control during systems operations. In addition, there are a number of key technology concerns, such as significant landing modes which tend to be closely spaced and distributed widely, distributed sensors and actuators which are collecated, and the wide distribution of landing modes that must be reduced from a dimensional standpoint.
An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model of the Douglas XF4D-1 airplane has been made in the Langley free-flight tunnel. The model was flown with leading-edge slats retracted and extended over a lift-coefficient range from 0.5 to the stall. Only relatively low-altitude conditions were simulated and no attempt was made to determine the effect on the stability characteristics of freeing the controls. The longitudinal stability and control characteristics of the model were satisfactory for all conditions investigated except near the stall with slats extended, where the model had a slight nosing-up tendency. The lateral stability and control characteristics of the model were considered satisfactory for all conditions investigated except near the stall with slats retracted, where a change in sign of the static- directional-stability parameter Cn(sub beta) caused the model to be directionally divergent. The addition of an extension to the top of the vertical tail did not increase Cn(sub beta) enough to eliminate the directional divergence of the model, but a large increase in Cn(sub beta) that was obtainable by artificial means appeared to eliminate the divergence and flights near the stall could be made. Artificially increasing the stability derivative-Cn(sub r) (yawing moment due to yawing) and Cn(sub p) (yawing moment due to rolling) had little effect on the divergence for the range of these parameters investigated. Calculations indicate that the damping of the lateral oscillation of the airplane with slats retracted or extended will be satisfactory at sea level but will be only marginally satisfactory at 40,000 feet.
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A study is made of aerodynamic performance and static stability and control at hypersonic speeds. In a first part of the study, the effect of interference lift is investigated by tests of asymmetric models having conical fuselages and arrow plan-form wings. The fuselage of the asymmetric model is located entirely beneath the wing and has a semicircular cross section. The fuselage of the symmetric model was centrally located and has a circular cross section. Results are obtained for Mach numbers from 3 to 12 in part by application of the hypersonic similarity rule. These results show a maximum effect of interference on lift-drag ratio occurring at Mach number of 5, the Mach number at which the asymmetric model was designed to exploit favorable lift interference. At this Mach number, the asymmetric model is indicated to have a lift-drag ratio 11 percent higher than the symmetric model and 15 percent higher than the asymmetric model when inverted. These differences decrease to a few percent at a Mach number of 12. In the course of this part of the study, the accuracy to the hypersonic similarity rule applied to wing-body combinations is demonstrated with experimental results. These results indicate that the rule may prove useful for determining the aerodynamic characteristics of slender configurations at Mach numbers higher than those for which test equipment is really available. In a second part of the study, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail. Results for Mach numbers from 3 to 18 for performance and 0.6 to 12 for stability and control are obtained by standard text techniques, by application of the hypersonic stability rule, and/or by use of helium as a test medium. Lift-drag ratios of about 5 for Mach numbers up to 18 are shown to be obtainable. The glider studied is shown to have acceptable longitudinal and directional stability characteristics through the range of Mach numbers studied. Some roll instability (negative effective dihedral) is found at Mach numbers near 12.
An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model with the outer wing panels swept 75 deg. has been conducted in the Langley 16-foot transonic tunnel. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. The engine nacelles incorporated swept lateral and vertical fins for aerodynamic stability and control. Jet-off data were obtained with flow-through nacelles, simulating inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained at Mach numbers from 0.60 to 1.05 through a range of angles of attack and angles of side-slip. Control characteristics were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control. The results indicate that the basic wing-body configuration becomes neutrally stable or unstable at a lift coefficient of 0.15; addition of nacelles with fins delayed instability to a lift coefficient of 0.30. Addition of nacelles to the wing-body configuration increased minimum drag from 0.0058 to 0.0100 at a Mach number of 0.60 and from 0.0080 to 0.0190 at a Mach number of 1.05 with corresponding reductions in maximum lift-drag ratio of 12 percent and 33 percent, respectively. The nacelle-fin combinations were ineffective as longitudinal controls but were adequate as directional and lateral controls. The model with nacelles and fins was directionally and laterally stable; the stability generally increased with increasing lift. Jet interference effects on stability and control characteristics were small but the adverse effects on drag were greater than would be expected for isolated nacelles.
Apollo Telescope Mount stabilization and control, with control moment gyro and experiment pointing control subsystems
An approach for designing inputs to identify stability and control derivatives from flight test data is presented. This approach is based on finding inputs which provide the maximum possible accuracy of derivative estimates. Two techniques of input specification are implemented for this objective - a time domain technique and a frequency domain technique. The time domain technique gives the control input time history and can be used for any allowable duration of test maneuver, including those where data lengths can only be of short duration. The frequency domain technique specifies the input frequency spectrum, and is best applied for tests where extended data lengths, much longer than the time constants of the modes of interest, are possible. These technqiues are used to design inputs to identify parameters in longitudinal and lateral linear models of conventional aircraft. The constraints of aircraft response limits, such as on structural loads, are realized indirectly through a total energy constraint on the input. Tests with simulated data and theoretical predictions show that the new approaches give input signals which can provide more accurate parameter estimates than can conventional inputs of the same total energy. Results obtained indicate that the approach has been brought to the point where it should be used on flight tests for further evaluation.
Stability and control derivative estimates obtained from the first three Space Shuttle entries are presented. The derivative estimates were obtained using the established modified maximum likelihood estimation program (MMLE3). The method of analysis used by the MMLE3 program is reviewed, the Shuttle configuration and data system are described, and the various types of maneuvers analyzed from the three entries are illustrated. Finally, the paper presents selected derivative results and compares them with predictions. Most of the flight-derived estimates agreed fairly well with predictions, considering the lack of experience in these new flight regimes. The most notable exception was the aerodynamic interference caused by firing the reaction control jets in the atmosphere. The flight results showed this interference to be considerably smaller than predicted.
A unique approach has been developed for stability and control derivative testing of the Space Shuttle orbiter during entry. Shuttle Program requirements have necessitated a minimum of testing. Therefore, flight tests concentrate on potential control problem areas predicted from wind tunnel data as well as anomalies discovered during flight testing. The object is to use the test data to remove center of gravity (cg), angle of attack, and elevon placards which are a function of these potential control problems. To ensure successful aerodynamic extraction from a minimum of testing, the following special measures have been taken. Exact maneuvers are designed preflight on Shuttle simulators. These maneuvers are duplicated and implemented during the flight by onboard software. An onboard instrumentation system was designed especially for aerodynamic parameter identification. State-of-the-art techniques are used in extracting aerodynamics.
The stability and control characteristics of an airship equipped with lifting rotors to provide a modest VTOL capability are discussed. The rotors are used for control and maneuvering in near-hovering flight. Configurations with two, three, and four lifting rotors are examined and compared with respect to control capabilities and dynamic response characteristics. Linearized models of the dynamics are employed for this study. A new approach to the prediction of rotor derivatives for operation near zero thrust in hover is presented. It is found that all three configurations have similar control and response characteristics. The responses are characterized by long time constants and low levels of angular damping.
Experimental aerodynamic investigations were conducted on a stingmounted 0.0405-scale representation (model 43-0) of the 140A/B Space Shuttle Orbiter in a Low Speed Wind Tunnel. The NASA designation for this test was 0A62A. The primary test objective was to continue studies, initiated on tests 0A16 and 0A71A and 0A71C, in optimizing the air breathing propulsion system (ABPS) and investigating the aerodynamic effects of various nacelle number/location configurations on the orbiter stability and control characteristics. Orbiter stability and control characteristics, both with and without ABPS, were investigated at elevon deflections of 0, + or -5, + or -19, + or -5, and -20 deg; aileron deflections of 0 and 10 deg (about 0 deg elevon); and rudder deflections of 0, -7.5, and -15 deg. Aerodynamic force and moment data was measured in the body axis system by a 2.5-inch task type internal balance. The model was sting supported through the base region with a nominal angle of attack range of -4 to 30 deg. Yaw polars were recorded over the beta range of -10 to 10 deg at fixed angles of attack of 0, 5, 10, and 15 deg.
This report contains the flight-test results of the lateral and directional-stability and control phase (including tests with wing-tip tanks) of a general flying-qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). These tests were conducted at indicated airspeeds up to 494 miles per hour (0.691 Mach number) at low altitude and up to 378 miles per hour (0.816 Mach number) at high altitude. These tests showed that the flying qualities of the airplane were for the most part in accordance with the requirements of the Army Air Forces Stability and Control Specifications. The only major deficiency noted was the negative lateral stability with the wing-tip tanks installed.
This Symposium is intended to bring together the often distinct cultures of the Stability and Control (S&C) community and the Computational Fluid Dynamics (CFD) community. The COMSAC program is itself a new effort by NASA Langley to accelerate the application of high end CFD methodologies to the demanding job of predicting stability and control characteristics of aircraft. This talk is intended to set the stage for needing a program like COMSAC. It is not intended to give details of the program itself. The topics include: 1) S&C Challenges; 2) Aero prediction methodology; 3) CFD applications; 4) NASA COMSAC planning; 5) Objectives of symposium; and 6) Closing remarks.
Tests were made of a model representative of a single-engine tractor-type airplane for the purpose of determining the stability and control effects of a propeller used as an aerodynamic brake. The tests were made with single-and dual-rotation propellers to show the effect of type of propeller rotation, and with positive thrust to provide basic data with which to compare the effects of negative thrust. Four configurations of the model were used to give the effects of tilting the propeller thrust axis down 5 deg., raising the horizontal tail, and combining both tilt and raised tail. Results of the tests are reported herein. The effects of negative thrust were found to be significant. The longitudinal stability was increased because of the loss of wing lift and increase of the angle of attack of the tail. Directional stability and both longitudinal and directional control were decreased because of the reduced velocity at the tail. These effects are moderate for moderate braking but become pronounced with full-power braking, particularly at high values of lift coefficient. The effects of model configuration changes were small when compared with the over-all effects of negative-thrust operation; however, improved stability and control characteristics were exhibited by the model with the tilted thrust axis. Raising the horizontal tail improved the longitudinal characteristics, but was detrimental to directional characteristics. The use of dual-rotation propeller reduced the directional trim charges resulting from the braking operation. A prototype airplane was assumed and handling qualities were computed and analyzed for normal (positive thrust) and braking operation with full and partial power. The results of these analyses are presented for the longitudinal characteristics in steady and accelerated flight, and for the directional characteristics in high- and low-speed flight. It was found that by limiting the power output of the engine (assuming the constant-speed propeller will function in the range of blade angles required for negative thrust) the stability and control characteristics may be held within the limits required for safe operation. Braking with full power, particularly at low speeds, is dangerous, but braking with very small power output is satisfactory from the standpoint of control. The amount of braking produced with zero power output is equal to or better than that produced by conventional spoiler-type brakes.
Estimates of the static stick-fixed stability and control characteristics of the Consolidated Vultee model 240 airplane are presented in this report. The estimates are based on tests of a 0.092-scale powered model in the 10-foot wind tunnel of the Guggenheim Aeronautical Laboratory of the California Institute of Technology. Results of the analysis are evaluated in terms of the Army specifications for stability and control characteristics which are more specific and, in general, equal to or more rigid than the Civil Aeronautics Administration requirements. The stick-fixed stability and control characteristics of the Consolidated Vultee model 240 were found to be satisfactory except for the following: 1) Marginal longitudinal stability in the landing approach (flaps 30 deg, 50% minimum continuous power) with aft center of gravity (31% M.A.C.); 2) Marginal rudder control to hold zero sideslip in a climb after take-off with asymmetric power (flaps 30 deg, left engine inoperative, right engine delivering take-off power) with maximum rudder throw limited to +/- 18 deg; 3) Marginal dihedral effect with flaps 40 deg and engines delivering maximum continuous power.
The effects of mass distribution on lateral stability and control characteristics of an airplane have been determined by flight tests of a model in the NACA free-flight tunnel. In the investigation, the rolling and yawing movements of inertia were increased from normal values to values up to five times normal. For each moment-of-inertia condition, combinations of dihedral and vertical-tail area representing a variety of airplane configurations were tested. The results of the flight tests of the model were correlated with calculated stability and control characteristics and, in general, good agreement was obtained. The tests showed the following effects of increased rolling and yawing moments of inertia: no appreciable change in spiral stability; reductions in oscillatory stability that were serious at high values of dihedral; a reduction in the sensitivity of the model to gust disturbances; and a reduction in rolling acceleration provided by the ailerons, which caused a marked increase in time to reach a given angle of bank. The general flight behavior of the model became worse with increasing moments of inertia but, with combinations of small effective dihedral and large vertical-tail area, satisfactory flight characteristics were obtained at all moment-of-inertia conditions.