Body motion and angles of attack during Project Fire Flight II reentry
Body motion and angles of attack during Fire Project flight reentry
ATMOSPHERIC ENTRY↗
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Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack
Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.
Boundary Layer Transition↗
An experimental study of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack
An experimental study of the effects of aft blowing on the forebody vortex asymmetry over a 3.0 caliber tangent ogive body at high angles of attack was conducted. The tip of the ogive body was equipped with a single blowing nozzle whose position could be adjusted. The tests were conducted in a subsonic wind tunnel at laminar flow conditions. The effects of model roll, angle of attack, blowing coefficient, and blowing nozzle axial position were independently studied. Surface pressure measurements and flow visualization results were obtained. Aft blowing was observed to alleviate the degree of vortex asymmetry at all angles of attack. The blowing was found to be more effective at the higher angles of attack. However, proportional control of the degree of vortex asymmetry was not observed, because the initial flowfield was highly asymmetric.
Gittner, Nathan M.↗
Development of a pneumatic high-angle-of-attack flush airdata sensing (HI-FADS) system
A nonintrusive high-angle-of-attack flush airdata sensing system was installed and flight tested in the F-18 High Alpha Research Vehicle. This system consists of a matrix of 25 pressure orifices arranged in concentric circles on the nose of the vehicle to determine angles of attack and sideslip, Mach number, and pressure altitude. During the course of the flight tests, it was determined that satisfactory results could be achieved using a subset of just nine ports.
Whitmore, Stephen A.↗
Turbulent viscous shock layer solutions for Jovian entry at small angles of attack
The equations governing the laminar and turbulent flows of reacting gas mixtures in chemical equilibrium over axially symmetric blunt bodies at small angles of attack are developed and presented in the unsteady conservative form. Solutions are obtained in the planes of symmetry of the flow field for the typical conditions encountered by a probe entering the Jovian atmosphere at small angles of attack. The eddy-viscosity is approximated by a two-layer model. The shock wave is treated as a discontinuity across which the shock relations are used to compute the flow conditions behind the shock. A time-asymptotic finite-difference method is used to solve the equations. The zero angle of attack results for a hyperboloid are compared with the existing results but neither experimental nor computational results are available for comparison with the present results at angle of attack.
Jumar, A.↗
Preliminary results from a subsonic high-angle-of-attack flush airdata sensing (HI-FADS) system - Design, calibration, algorithm development, and flight test evaluation
A nonintrusive high angle-of-attack flush airdata sensing (HI-FADS) system was installed and flight-tested on the F-18 high alpha research vehicle. This paper discusses the airdata algorithm development and composite results expressed as airdata parameter estimates and describes the HI-FADS system hardware, calibration techniques, and algorithm development. An independent empirical verification was performed over a large portion of the subsonic flight envelope. Test points were obtained for Mach numbers from 0.15 to 0.94 and angles of attack from -8.0 to 55.0 deg. Angles of sideslip ranged from -15.0 to 15.0 deg, and test altitudes ranged from 18,000 to 40,000 ft. The HI-FADS system gave excellent results over the entire subsonic Mach number range up to 55 deg angle of attack. The internal pneumatic frequency response of the system is accurate to beyond 10 Hz.
Whitmore, Stephen A.↗
Internal flow calculations for axisymmetric supersonic inlets at angle of attack
Several solutions for the internal flow in axisymmetric supersonic inlets at angle of attack were obtained. A shock-capturing finite-difference technique was used to obtain the solutions. Good overall agreement between the shock-capturing solutions and experimental data was obtained, except in regions of strong viscous effects or boundary-layer removal. The required centerbody translation with changes in freestream Mach number and angle of attack was obtained and agrees qualitatively with experimental data.
Presley, L. L.↗
Flight-determined correction terms for angle of attack and sideslip
The effects of local flow, upwash, and sidewash on angle of attack and sideslip (measured with boom-mounted vanes) were determined for subsonic, transonic, and supersonic flight using a maximum likelihood estimator. The correction terms accounting for these effects were determined using a series of maneuvers flown at a large number of flight conditions in both augmented and unaugmented control modes. The correction terms provide improved angle-of-attack and sideslip values for use in the estimation of stability and control derivatives. In addition to detailing the procedure used to determine these correction terms, this paper discusses various effects, such as those related to Mach number, on the correction terms. The use of maneuvers flown in augmented and unaugmented control modes is also discussed.
Shafer, M. F.↗
Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack
Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.
boundary layer transition↗
Wing-alone aerodynamic characteristics for high angles of attack of supersonic speeds
An experiment was conducted to determine wing-alone supersonic aerodynamic characteristics at high angles of attack. The wings tested varied in aspect ratio from 0.5 to 4.0 and in taper ratio from 0 to 1.0. The wings were tested at angles of attack ranging rom -5 deg to 60 deg and at Mach number from 1.60 to 4.60. The aerodynamic characteristics were obtained by integrating local pressures measured over the wing surfaces. Presented and discussed are results showing the effects of aspect ratio, taper ratio, Mach number, and angle of attack on force and moment coefficients and center of pressure locations. Also included are tabulations of the pressure measurements.
Stallings, R. L., Jr.↗
Analysis of supersonic-hypersonic flutter of lifting surfaces at angle of attack
Lifting surfaces supersonic-hypersonic flutter at angle of attack determined by shock expansion, Newtonian flow and local flow piston theory
Bennett, R. M.↗
Aircraft aerodynamics at high angles of attack
AEROX program estimates aircraft aerodynamics to high angles of attack (up to sixty degrees). It estimates coefficients of lift, induced drag, and pitching moment for wings and wing body combinations with or without aft horizontal tail. Both trimmed and untrimmed characters are calculated.
Axelson, J. A.↗
Analysis of the longitudinal handling qualities and pilot-induced-oscillation tendencies of the High-Angle-of-Attack Research Vehicle (HARV)
The NASA High-Angle-of Attack Research Vehicle (HARV), a modified F-18 aircraft, experienced handling qualities problems in recent flight tests at NASA Dryden Research Center. Foremost in these problems was the tendency of the pilot-aircraft system to exhibit a potentially dangerous phenomenon known as a pilot-induced oscillation (PIO). When they occur, PIO's can severely restrict performance, sharply dimish mission capabilities, and can even result in aircraft loss. A pilot/vehicle analysis was undertaken with the goal of reducing these PIO tendencies and improving the overall vehicle handling qualities with as few changes as possible to the existing feedback/feedforward flight control laws. Utilizing a pair of analytical pilot models developed by the author, a pilot/vehicle analysis of the existing longitudinal flight control system was undertaken. The analysis included prediction of overall handling qualities levels and PIO susceptability. The analysis indicated that improvement in the flight control system was warranted and led to the formulation of a simple control stick command shaping filter. Analysis of the pilot/vehicle system with the shaping filter indicated significant improvements in handling qualities and PIO tendencies could be achieved. A non-real time simulation of the modified control system was undertaken with a realistic, nonlinear model of the current HARV. Special emphasis was placed upon those details of the command filter implementation which could effect safety of flight. The modified system is currently awaiting evaluation in the real-time, pilot-in-the-loop, Dual-Maneuvering-Simulator (DMS) facility at Langley.
Hess, Ronald A.↗
Heating and flow-field studies on a straight-wing hypersonic reentry vehicle at angles of attack from 20 to 80 deg with simulation of real-gas trends
Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.
Hunt, J. L.↗
Flight angle of attack determined from ablative heat-shield patterns.
Hypersonic trim angle of attack of lifting entry vehicles, correlating stagnation point values obtained from heat shield ablation and pressure distribution measurements
PRESSURE DISTRIBUTION↗
Study of flow fields about axisymmetric blunt bodies at large angle-of-attack Final report
Flow fields about axisymmetric blunt bodies at large angles of attack calculated by IBM 7094 computer program
FLOW FIELD↗
Evaluation of five conical center-body supersonic diffusers at several angles of attack
Five supersonic inlets having conical center bodies were investigated in a 16-inch ram jet engine with and without combustion at Mach numbers of 1.7, 1.9, and 2.0 and angles of attack from 0 degrees to 10 degrees. Four of the inlets were of the low-mass-flow-ratio type whereas the fifth inlet was of the high-mass-flow-ratio type. The stable operating range of the four low-nass-flow-ratio inlets decreased considerably as the angle of attack was increased despite an increase of the cone angle or the use of boundary-layer bleed. The stable range of the high-mass-flow-ratio inlet increased considerably both at 0 degrees and at greater angles of attack as the free-stream Mach number was decreased. Free-stream Mach number, angle of attack, and inlet type had negligible effects on combustion efficiency, temperature ratio, or combustion-chamber inlet Mach number in the supercritical region.
Englert, Gerald W↗
High angle of attack hypersonic aerodynamics
A new aerodynamics force model is presented which is based on modified Newtonian theory and empirical correlations. The algebraic model was developed for complete vehicles from take off to orbital speeds and for large angles of attack. Predictions are compared to results for a wind tunnel model at a Mach number of 20, and the full scale Shuttle Orbiter for Mach numbers from 0.25 to 20 for angles of attack from 0 to 50 deg. The maximum shuttle orbiter lift/drag at Mach 10 and 20 is 1.85 at 20-deg angle-of-attack. Aerodynamic force predictions are made for a transatmospheric vehicle, which is a derivative of the Shuttle Orbiter, for Mach numbers from 4 to 27 at angles of attack from 5 to 40 deg. Predicted aerodynamic force data indicate that lift/drag ratios of 5.2 at Mach number 10 and 3.6 at Mach number 26 are obtainable. Changes in force coefficients with changes in: nose angle, sweep angle, and (volume exp 2/3)/planform area are quantified for Mach numbers of 10 and 26. Lift/drag ratios increase with decreasing nose angle and (volume exp 2/3)/planform area and increasing wing sweep angle. Lift/drag ratios are independent of these variables for angles of attack in excess of 20 deg at Mach 10 and 30 deg at Mach 26.
Harloff, Gary J.↗