Engineering topics
Publications and source records attributed to Ashby, G. C., Jr..
Flow survey instrumentation integral part of model support system. Drive motor, limit switches, and position transducer contained within streamlined housing and operable in near vacuum wing-tunnel environment. Sting-mounted system has advantages over conventional wall-mounted flow-field survey equipment, include more efficiently utilized run time, higher position accuracy, and fewer runs to map flow field.
The procedure described involves calibrating the probe in the gas of interest over the range of temperatures and pressures anticipated in the wind-tunnel tests and then applying the calibration to reduce the data from those tests. Pressure readings from two probes, one having two hot-film sensors and an internal sonic orifice and the other connected to a conventional pressure gage, are recorded for several Reynolds numbers and compared. The conventional static-pressure probe with a pair of hot-film sensors and an interior sonic orifice is shown to provide the means of surveying flowfields in hypersonic flow at a rapid rate and with an accuracy of better than + or - 10%.
A static pressure probe employing hot-film sensors has been developed for the rapid measurement of the static pressure fields surrounding analytic shapes in hypersonic flows. The hot-film probe is a modification of the standard static pressure probe, consisting of a front hot-film sensor operated as a resistance thermometer, a rear sensor operated at an overheat ratio of 1.5 to 1.8 and a small sonic orifice installed inside the tubing of a conventional device. The probe has been calibrated in helium and air over a range of temperatures and pressures in a bell jar apparatus, with a repeatability of the data to within + or - 0.015 mm Hg. Comparative tests of the hot-film and conventional static pressure probes in a hypersonic helium wind tunnel at Mach 20 and various Reynolds numbers have indicated the settling time of the hot-film probe to be on the order of milliseconds, as compared with 30 sec for the conventional probe. The pressures measured by the two probes were found to be within 10% of each other. Although the hot-film probe makes flow-field static pressure surveys more practical in blowdown hypersonic wind tunnels, viscous and flow angle effects still must be assessed under the conditions of use.
An experimental and analytical study of the sonic boom static pressure signatures generated by a blunt body at mach 6 has shown that finite difference computer programs can be used to give reasonable estimates of the signatures. The calculated near field static pressure signature was extrapolated to the far field by a program using the method of characteristics. A comparison of this extrapolated signature with the signature predicted by far field sonic boom theory (linearized) shows that peak overpressures are about the same, at least up to mach, but the far field theory overestimates the length of the signature.
Static-pressure signatures parallel to the flight path of the launch and entry configurations of the space shuttle were measured. The launch configurations, consisting of an equivalent body of revolution (representing the orbiter and external fuel tank) with a solid exhaust gas plume attached, was tested at an angle of attack of 0 deg. The entry configuration (orbiter only) was tested over an angle-of-attack range from 10 deg to 40 deg. Calculated pressure signatures were in good agreement with measured signatures for both configurations.
Pitot pressure and flow angle distributions in the windward flow field of the NASA 040A space shuttle orbiter configuration and surface pressures were measured, at a Mach number of 20 and an angle of attack of 31 deg. The free stream Reynolds number, based on model length, was 5.39 x 10 to the 6th power. Results show that cores of high pitot pressure, which are related to the body-shock-wing-shock intersections, occur on the windward plane of symmetry in the vicinity of the wing-body junction and near midspan on the wing. Theoretical estimates of the flow field pitot pressures show that conical flow values for the windward plane of symmetry surface are representative of the average level over the entire lower surface.
A testing technique for obtaining the static aerodynamic characteristics of twin-fuselage configurations at hypersonic speeds by using a conventional single-balance installation has been evaluated. Data from a triple-fuselage model and a single-fuselage model were summed and then halved to obtain the characteristics for a twin-fuselage model of the same scale. The three related models were evaluated experimentally at Mach 20.3 in helium and Mach 6 in air for an angle-of-attack range from minus 6 deg to 50 deg. The Reynolds numbers, based on model length, were 1.88 million for the Mach 20.3 tests and 2.55 million for the Mach 6 tests.
Wave and skin-friction drag have been numerically calculated for a series of power-law bodies at a Mach number of 6 and Reynolds numbers, based on body length, from 1.5 million to 9.5 million. Pressure distributions were computed on the nose by the inverse method and on the body by the method of characteristics. These pressure distributions and the measured locations of boundary-layer transition were used in a nonsimilar-boundary-layer program to determine viscous effects. A coupled iterative approach between the boundary-layer and pressure-distribution programs was used to account for boundary-layer displacement-thickness effects. The calculated-drag coefficients compared well with previously obtained experimental data.
Experimental data have been obtained for two series of bodies at Mach 6 and Reynolds numbers, based on model length, from 1.4 million to 9.5 million. One series consisted of axisymmetric power-law bodies geometrically constrained for constant length and base diameter with values of the exponent n of 0.25, 0.5, 0.6, 0.667, 0.75, and 1.0. The other series consisted of positively and negatively cambered bodies of polygonal cross section, each having a constant longitudinal area distribution conforming to that required for minimizing zero-lift wave drag at hypersonic speeds under the geometric constraints of given length and volume. At the highest Reynolds number, the power-law body for minimum drag is blunter (exponent n lower) than predicted by inviscid theory (n approximately 0.6 instead of n = 0.667); however, the peak value of lift-drag ratio occurs at n = 0.667. Viscous effects were present on the bodies of polygonal cross section but were less pronounced than those on the power-law bodies. The trapezoidal bodies with maximum width at the bottom were found to have the highest maximum lift-drag ratio and the lowest mimimum drag.
A one hundredth scale model of the modified 089B shuttle orbiter was tested in the Langley 20-Inch Mach 6 tunnel. Force and moment, surface pressure and oilflow data were obtained on one model, and phase-change coating data were obtained on another. The pressure tests were conducted first; the tubes were clipped near the base of the model and then the force and moment and oil flow tests conducted. Angles of attack for the tests were from 20 deg to 35 deg and are commensurate with the range of flight values from entry down to Mach 5. The design flight Reynolds number at Mach 6, based on model length, was 15 million, which could not be obtained in the tunnel; therefore, the tests were conducted at the highest and lowest values for this model in the tunnel, 9.4 million and 4.0 million, respectively, to indicate Reynolds number effects. Two control deflection combinations, representative of the bank and pitch control limits of the design flight trajectory, were used.
A scale model of the North American Rockwell ATP Orbiter with and without the external tank has been tested in a 22-inch helium tunnel at Mach 20 and a Reynolds number based on model length, of 2.14 times one million. Longitudinal and lateral-directional data were determined for the orbiter alone while only longitudinal characteristics and elevon roll effectiveness were investigated for the orbiter/tank combination. Oil flow and electron beam flow visualization studies were conducted for the orbiter alone, orbiter with external tank and the ascent configuration.
The longitudinal, lateral, and directional aerodynamic characteristics of a delta-wing configuration were obtained experimentally at Mach 20 in helium with Reynolds numbers, based on model length, of 1.5 million and 2.9 million and at a Mach number of 6 in air with a Reynolds number, based on model length, of 4.8 million. The angles of attack varied from 0 deg to 55 deg for two sideslip angles. The effects of the addition of dorsal fins, the removal of wing tip fins, an increase in elevon span, and changes in elevon hinge-line sweep angle are discussed. The unmodified vehicle had a maximum lift-drag ratio of 2.1 at Mach 19 and of 2.4 at Mach 6 with about the same lateral and directional stability level at both Mach numbers. As the Mach number increased from 6 to 20, the longitudinal center of pressure moved forward and more positive elevon deflection was therefore required to maintain a given trim angle. The removal of wing tip fins increased the maximum lift-drag ratio and had a negligible effect on longitudinal stability, but caused directional instability that was not corrected by the dorsal fins examined. The shape of the wing and elevon hinge-line sweep had a large influence on the induced yawing moment due to roll control.
Pitot-pressure profiles and surface pressure measurements have been obtained at four axial stations along the center lines of space shuttle straight-wing and delta-wing orbiters at angles of attack from 20 deg to approximately 60 deg and Mach numbers of 20.3 (helium) and 6.8 (air). The results show that the Mach number at the edge of the boundary layer is best predicted by tangent-cone theory up to shock detachment.
Comparisons of hinge moments for simple delta wing and delta wing orbiter concept at Mach 6
Hypersonic body shaping for minimum drag and improved performance from view of flight regimes, discussing implications of pressure laws
Aerodynamic characteristics of modified conical lifting reentry vehicle at hypersonic speeds
Volume addition effects on leeward surface of high lift drag ratio winged body at hypersonic speeds
Calculated and experimental hinge moments on trailing-edge flaps of swept delta wings at hypersonic speeds and positive deflection angles