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NASA NTRS1989

A numerical analysis of the British Experimental Rotor Program blade

Two Computational Fluid Dynamic codes which solve the compressible full-potential and the Reynolds-Averaged Thin-Layer Navier-Stokes equations were used to analyze the nonrotating aerodynamic characteristics of the British Experimental Rotor Program (BERP) helicopter blade at three flow regimes: low angle of attack, high angle of attack and transonic. Excellent agreement was found between the numerical results and experiment. In the low angle of attack regime, the BERP had less induced drag than a comparable aspect ratio rectangular planform wing. At high angle of attack, the blade attained high-lift by maintaining attached flow at the outermost spanwise locations. In the transonic regime, the BERP design reduces the shock strength at the outer spanwise locations which affects wave drag and shock-induced separation. Overall, the BERP blade exhibited many favorable aerodynamic characteristics in comparison to conventional helicopter rotor blades.

Duque, Earl P. N.↗

NASA NTRS1990

Decomposition technique and optimal trajectories for the aeroassisted flight experiment

An actual geosynchronous Earth orbit-to-low Earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking Earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.

Miele, A.↗

NASA NTRS1991

Decomposition technique and optimal trajectories for the aeroassisted flight experiment

An actual geosynchronous earth orbit-to-low earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.

Miele, A.↗

NASA NTRS1958

Wind-Tunnel Investigation of Some Effects of Wing Sweep and Horizontal-Tail Height on the Static Stability of an Airplane Model at Transonic Speeds

A research model of an airplane with a configuration suitable for supersonic flight was tested at transonic speeds in order to establish the effects on longitudinal and lateral stability of certain changes in both wing sweep and height of the horizontal tail. Two wings of aspect ratio 3 and taper ratio 0.15, one having the quarter-chord line swept back 30 deg and the other 45 deg, were each tested with the horizontal tail of the model in a low and in a high position. One configuration was also tested with fuselage strakes. The tests were made at Mach numbers from 0.60 to 1.17 and Reynolds numbers from 1.9 x 10(exp 6) to 2.6 x 10(exp 6). The results indicated that a low horizontal-tail position (below the wing-chord plane) gave positive longitudinal stability for the model for all angles of attack used (angles of attack up to 24 deg); whereas, a higher tail position (above the wing-chord plane) resulted in a large reduction in stability at moderate angles of attack. With the higher horizontal tail, the 30 deg-swept-wing model had somewhat more stability than the 45 deg-swept-wing model at subsonic Mach numbers. With the lower tail, the 45 deg-swept-wing model had slightly more stability at all Mach numbers. The model with the 30 deg swept wing had greater directional stability with the tail in the higher rather than the lower position, but the opposite was true for the 45 deg-swept-wing model. The directional stability decreased sharply at high angles of attack; this characteristic was alleviated by the use of fuselage strakes which, however, proved to be detrimental to the longitudinal stability of the model tested.

Fisher, Lewis R.↗

NASA NTRS1954

Stability and Control Flight Tests of a Vertically Rising Airplane Model Similar to the Lockheed XFV-1 Airplane

This paper presents the results of an investigation of the dynamic stability and controllability of a model which approximately represents the Lockheed XFV-1 airplane to a 1/8 scale. The investigation consisted of hovering flights in still air at a considerable height above the ground, hovering flights very close to the ground, vertical take-offs and landings, flights through the transition range from hovering to normal forward flight, and sideways translational flights. The model could be flown smoothly and easily in hovering flight despite the fact that the uncontrolled pitching and yawing motions were unstable oscillations. There was a noticeable reduction in the controllability of the model when hovered very close to the ground but take-offs could be made easily and landings on a g,ven spot could be made accurately in spite of this adverse ground effect. Flights through the transition range from hovering to normal forward flight could be performed fairly easily. The model seemed to have stability of angle of attack and angle of roll over most of the transition range. The yawing motion was divergent in the very high angle-of-attack range but could be controlled easily. At the lower angles of attack, the model seemed to become stable in yaw. In sideways flight there was an increasingly strong tendency to diverge in roll as the speed was increased and finally, at a speed of about 25 knots (full scale), the model rolled off despite efforts of the pilot to control it.

Kirby, Robert H.↗

NASA NTRS1997

Analysis of Wind Tunnel Longitudinal Static and Oscillatory Data of the F-16XL Aircraft

Static and oscillatory wind tunnel data are presented for a 10-percent-scale model of an F-16XL aircraft. Static data include the effect of angle of attack, sideslip angle, and control surface deflections on aerodynamic coefficients. Dynamic data from small-amplitude oscillatory tests are presented at nominal values of angle of attack between 20 and 60 degrees. Model oscillations were performed at five frequencies from 0.6 to 2.9 Hz and one amplitude of 5 degrees. A simple harmonic analysis of the oscillatory data provided Fourier coefficients associated with the in-phase and out-of-phase components of the aerodynamic coefficients. A strong dependence of the oscillatory data on frequency led to the development of models with unsteady terms in the form of indicial functions. Two models expressing the variation of the in-phase and out-of-phase components with angle of attack and frequency were proposed and their parameters estimated from measured data.

Klein, Vladislav↗

NASA NTRS1981

Estimation of dynamic stability parameters from drop model flight tests

The overall remotely piloted drop model operation, descriptions, instrumentation, launch and recovery operations, piloting concept, and parameter identification methods are discussed. Static and dynamic stability derivatives were obtained for an angle attack range from -20 deg to 53 deg. It is indicated that the variations of the estimates with angle of attack are consistent for most of the static derivatives, and the effects of configuration modifications to the model were apparent in the static derivative estimates.

Chambers, J. R.↗

NASA NTRS1983

Acoustic measurements of the X-wing rotor

Noise measurements of a stoppable X-wing rotor system model, tested in the Ames 40- by 80-foot wind tunnel, are summarized. Performance, control system stability, and noise of the model were investigated at various forward speeds, tip speeds, collective blade angles, jet blowing velocities, and model attack angles. The model was tested in the rotating wing helicopter configuration, in the fixed wing configuration, and in wing configurations between the two. Noise data obtained in the helicopter configuration at the two highest tip speeds (Mach 0.44 and 0.47) and at wind tunnel speeds below 140 knots are reported. Test configuration and performance information are included. General acoustic measurements (dB, dBA, and PNdB) at six microphone locations are presented for all conditions under which the background noise was below the model noise. More specific measurements (1/3-octave and blade passage frequency harmonic levels) are presented for selected conditions. Graphs of dBA and 1/3-octave spectra, which show the noise trends as functions of operating condition, are included. The noise depends mainly on the jet blowing velocity. The noise levels were highest at moderate jet blowing velocities, less at the highest velocity, and lowest with no blowing at all.

Mosher, M.↗

NASA NTRS1971

Heat transfer tests of the NASA-MSC space shuttle configuration at the Langley Research Center Mach 8 Variable Density Facility

The experimental investigations performed on the NASA-Manned Spacecraft Center Space Shuttle orbiter and booster configurations at a Mach 8 variable density facility are presented. The test program was a series of aerothermodynamic wind tunnel tests that were run over a range of angles of attack, yaw angles, and Reynolds numbers. Objectives of the test program were to obtain heat transfer data over the NASA-Manned Spacecraft Center Space Shuttle orbiter, booster, and launch configurations for a range of angles of attack from - 20 to + 30 deg, yaw angles of 0 and + or - 6 deg, and Reynolds numbers of 0.6, 2.0, and 3.7 x one million. The phase-change coating technique was used to obtain heat transfer data. Information received from these tests will be instrumental in performing thermal protection systems studies and vehicle aerodynamic design.

Connor, L. E.↗

NASA NTRS1990

Maximum orbit plane change with heat-transfer-rate considerations

Two aerodynamic maneuvers are considered for maximizing the plane change of a circular orbit: gliding flight with a maximum thrust segment to regain lost energy (aeroglide) and constant altitude cruise with the thrust being used to cancel the drag and maintain a high energy level (aerocruise). In both cases, the stagnation heating rate is limited. For aeroglide, the controls are the angle of attack, the bank angle, the time at which the burn begins, and the length of the burn. For aerocruise, the maneuver is divided into three segments: descent, cruise, and ascent. During descent the thrust is zero, and the controls are the angle of attack and the bank angle. During cruise, the only control is the assumed-constant angle of attack. During ascent, a maximum thrust segment is used to restore lost energy, and the controls are the angle of attack and bank angle. The optimization problems are solved with a nonlinear programming code known as GRG2. Numerical results for the Maneuverable Re-entry Research Vehicle with a heating-rate limit of 100 Btu/ft(2)-s show that aerocruise gives a maximum plane change of 2 deg, which is only 1 deg larger than that of aeroglide. On the other hand, even though aerocruise requires two thrust levels, the cruise characteristics of constant altitude, velocity, thrust, and angle of attack are easy to control.

Lee, J. Y.↗

NASA NTRS1932

On the theory of wing sections with particular reference to the lift distribution

This report gives a simple and exact method of calculating the lift distribution on thin wing sections. The most essential feature of the new theory is the introduction of an "ideal angle of attack," this angle being defined as that at which the flow enters the leading edge smoothly or, more precisely, as the angle of attack at which the lift at the leading edge equals zero. The lift distribution at this particular angle is shown to be a characteristic property of the section and has been termed the "basic distribution." it is shown that the lift of a wing section may always be considered to consist of (a) the basic distribution and (b) the additional distribution, where the latter is independent of the mean camber line and thus identical for all thin sections. The specific reason for the poor aerodynamic qualities of thin wing sections is pointed out as being due to the fact that the additional lift in potential flow becomes infinite at the leading edge. It is established that the essential parameter occurring in this analysis is the radius of curvature at the leading edge. It is further established that the actual slope of the lift curve is dependent on this parameter. The theoretical lift distribution is compared with the distribution obtained by direct measurement on a number of the more conventional wing sections. The new theory is of value in calculating the strength of wing sections.

Theodorsen, Theodore↗

NASA NTRS2019

CHANGO: A Software Tool for Boost Stage Guidance of the Space Launch System Exploration Mission 1

The Day of Launch Initiation Load Update (DOLILU) System is the means by which the Space Launch System (SLS) Vehicle trajectory is designed, verified, and uploaded on the Day of Launch (DOL) in order to ensure a safe flight. Launch vehicles are designed to fly down a narrow angle of attack and sideslip angle corridor in order to keep them within structural load limits. The angle of attack and sideslip angle response to the launch vehicle experiences can vary significantly based upon the winds experienced on the DOL. SLS Boost Stage flight employs an open-loop guidance scheme through Solid Rocket Booster (SRB) separation. In the SLS open-loop scheme, the vehicle will fly a prescribed set of attitudes as a function of the change in altitude since launch. This set of reference attitude values and corresponding altitude reference independent values are designed with ground software using winds measured on the DOL with the goal of minimizing angle of attack and sideslip angle, thereby minimizing related ascent integrated vehicle structural loads. The table of Boost Stage attitude commands as a function of altitude gained since launch is called the chi table. A software tool called CHANGO (Chi Angle Optimizer) designs the Boost Stage chi table which is uploaded to the vehicle’s flight computer and used during ascent by the flight software (FSW). The wind and atmospheric conditions are measured prior to launch and pre-processed to become input to the CHANGO software along with a set of parameters developed in advance of the DOL. CHANGO’s target set consists of the heading and altitude rate at SRB separation determined well before launch by the Program to Optimize Simulated Trajectories (POST). CHANGO consists of a simplified three degree-of-freedom (3-DOF) simulation representing the SLS launch configuration. In general, the launch azimuth is strongly correlated with the heading at SRB separation, and the initial pitchover rate is strongly correlated with the altitude rate at SRB separation. CHANGO uses an adaptation of Powell’s method to vary the initial pitchover rate and launch azimuth to solve a 2-dimentional minimization problem. CHANGO’s trajectory simulation is phase-based, with flight events separating the phases. Each flight phase has different attitude alignment logic. CHANGO’s 3-DOF simulation starts when the vehicle’s thrust-to-weight ratio equals one, and ends at a pre-calculated SRB separation time.

Ahmad, Naeem↗

NASA NTRS1973

Space shuttle: Preliminary pressure distributions on the 049 orbiter, orbiter in presence of H/O tank and orbiter in launch configuration

The 049 orbiter and launch configurations were tested in a trisonic wind tunnel to obtain preliminary loads information on the orbiter alone, orbiter in presence of the H/O tank and orbiter in the full launch configuration. The orbiter consisted of the baseline 049 double-delta wing, twin vertical stabilizers, seven degrees of dihedral and included abort rockets. The orbiter was mounted at minus 1.50 degree of incidence (fuselage centerline relative to H/O tank centerline) in the launch configuration. The solid rocket motors were mounted at a radial location of 21 degrees from the horizontal centerline of the H/O tank. The test was conducted over a Mach number range of 0.6 to 4.96. Nominal angle of attack and angle of sideslip ranges of minus 6 to plus 6 degrees were tested. In addition, the orbiter alone was tested over an angle of attack range of minus 6 to plus 26 degrees.

Hamilton, J. T.↗

NASA NTRS1982

Aerodynamic interactions between a 1/6 scale helicopter rotor and a body of revolution

A wind-tunnel investigation was conducted in which independent, steady state aerodynamic forces and moments were measured on a 2.24-m-diam, two bladed helicopter rotor and a body of revolution. The objective was to determine the interaction of the body on the rotor performance and the effect of the rotor on the body aerodynamics for variations in velocity, thrust, tip-path-plane angle of attack, body angle of attack, rotor/body position, and body nose geometry. Results show that a body of revolution near the rotor can produce significant favorable or unfavorable effects on rotor performance, depending on the operating condition. Body longitudinal aerodynamic characteristics are significantly modified by the presence of an operating rotor and hub.

Betzina, M. D.↗

NASA NTRS1983

Interference effects of aft reaction-control yaw jets on the aerodynamic characteristics of a space shuttle orbiter model at supersonic speeds

A wind tunnel investigation of the interference effects of aft reaction control system yaw jet plumes on a 0.0125 scale Space Shuttle orbiter model was conducted at Mach numbers from 2.50 to 4.50. Test variables included model angle of attack, model angle of sideslip, jet to free stream mass flow ratio, and number and position of operating jets. The aft reaction control jet plume creates a blockage above and behind the wing on the side in which the jet exhausts and results in flow separation on the wing upper surface and fuselage side. Positive pitching moment and side force increments and negative yawing moment and rolling moment increments due to the flow separations are incurred for left side firing jets, primarily at angles of attack above 10 deg. The yawing moment interference increments are favorable and result in a small jet thrust amplification. As a result of this investigation, the aft reaction control system was certified for operation at supersonic Mach numbers prior to the first flight of the space transportation system (STS-1).

Covell, P. F.↗

NASA NTRS1983

Rotor/body aerodynamic interactions

A wind tunnel investigation was conducted in which independent, steady state aerodynamic forces and moments were measured on a 2.24 m diam. two bladed helicopter rotor and on several different bodies. The mutual interaction effects for variations in velocity, thrust, tip-path-plane angle of attack, body angle of attack, rotor/body position, and body geometry were determined. The results show that the body longitudinal aerodynamic characteristics are significantly affected by the presence of a rotor and hub, and that the hub interference may be a major part of such interaction. The effects of the body on the rotor performance are presented.

Betzina, M. D.↗

NASA NTRS1985

Wind-tunnel evaluation of a 21-percent-scale powered model of a prototype advanced scout helicopter

An exploratory wind tunnel investigation of a 21 percent scale powered model of a prototype advanced scout helicopter was conducted in the Langley 4 by 7 Meter Tunnel. The investigation was conducted to define the overall aerodynamic characteristics of the Army Helicopter Improvement Program (AHIP), to determine the effects of the rotor on the aerodynamic characteristics and to evaluate the effect of a mast mounted sight on the aircraft stability characteristics. Tests covered a range of thrust coefficients, advance ratios, angles of attack and angles of sideslip and were run for both rotor on and rotor off configurations. Results of the investigation showed that the prototype configuration was longitudinally unstable with angle of attack for all configurations tested. The instability was due to unfavorable interference effects between the horizontal tail and the wake shed from the engine pylon and rotor hub, which caused a loss of horizontal tail effectiveness. The addition of the mast mounted sight had little effect on the stability of the model, but it caused an alteration in the rotor lift distribution that resulted in substantial interference drag for the sight.

Phelps, A. E., III↗

NASA NTRS1985

Rotor/body aerodynamic interactions

Betzina, M. D.↗