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An investigation was conducted to design instruments to directly measure skin friction along the chamber walls of supersonic combustor models. Measurements were made in a combustor at the General Applied Science Laboratory (GASL) and in the Direct Connect Arcjet Facility (DCAF) supersonic combustor at the NASA AMES Research Center. Flow conditions in the high enthalpy combustor models ranged from total pressures of 275-800 psia (1900-5550 kPa) and total temperatures from 5800-8400 R (3222-4667 K). This gives enthalpies in the range of 1700-3300 BTU/Ib_m (3950-7660 KJ/kg) and simulated flight Mach number from 9 to 13. A direct force measurement device was used to measure the small tangential shear force resulting from the flow passing over a non-intrusive floating element. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the shear force on the order of 0.0005 in (0.0125 mm). This small deflection allows the balance to be a non-nulling type. Several measurements were conducted in cold supersonic flows to verify the concept and establish accuracy and repeatability. This balance design includes actively controlled cooling of the floating sensor head temperature through an internal cooling system to eliminate nonuniform temperature effects between the head and the surrounding chamber wall. This enabled the device to be suitable for shear force measurement in very hot flows. The key to this device is the use of a quartz tube cantilever with strain gages bonded at orthogonal positions directly on the surface at the base. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Bench tests showed that this did not influence the force measurement. Numerical heat transfer calculations were conducted for design feasibility and analysis, and to determine the effectiveness of the active cooling of the floating head. Analysis of the measurement uncertainty in cold supersonic flow tests show that uncertainty under 8% is achievable, but variations in the balance cooling during a particular test raised uncertainty up to 20% in these very hot flows during the early tests. Improvements to the strain gages and balance cooling reduced uncertainty for the later tests to under 15%. / Ph. D.

Experiments were performed with two different helium injector models at different injector transverse and yaw angles in order to determine the mixing rate and core penetration of the injectant and the flow field total pressure losses. when gaseous injection occurs into a supersonic freestream. Tested in the Virginia Tech supersonic tunnel. with a freestream Mach number of 3.0 and conditions corresponding to a freestream Reynolds number of 5.0 x 107 1m. was a single. sonic. 5X underexpanded, helium jet at a downstream angle of 30° relative to the freestream. This injector was rotated from 0° to _28° to test the effects of injector yaw. The second model was an array of three supersonic, 5X underexpanded helium injectors with an exit Mach number of 1.7 and a transverse angle of 15°. This model was tested in the NASA Langley Mach 6.0, High Reynolds number tunnel, with freestream conditions corresponding to a Reynolds number of 5.4 x 10⁷ /m. The injector array as tested at yaw angles of 0° and -15°. Surface flow visualization showed that significant flow asymmetries were produced by injector yaw. Nanosecond exposure shadowgraph pictures were taken, showing the gaseous injection plume to be unsteady, and further studies demonstrated this unsteadiness was related to shock waves orthogonal to the injectant bow shock, that were generated at a frequency of 30 kHz. The primary data technique used, was a concentration probe which measured the molar concentration of helium in the flow field. Concentration data and other meanflow data was taken at several downstream axial stations and yielded contours of helium concentration, total pressure, Mach number, velocity, and mass flux, as well as the static properties. From these contour plots, the various mixing rates for each case were determined. The injectant mixing rates, expressed as the maximum concentration decay, and mixing distances were found to be unaffected by injector yaw, in the Mach 3.0 experiments, but were adversely affected by injector yaw in the Mach 6.0 experiments. One promising aspect of injector yaw was the that as the yaw angle was increased, lateral motion of the injectant plume became significant, and the turbulent mixing region area increased by approximately 34%. Comparisons of the 15° transverse angled injection into a Mach 6.0 flow to previous experiments with 15° injection into a Mach 3.0 freestream, demonstrated that there is a significant decrease in initial mixing, at Mach 6.0, resulting in a much longer mixing distance. From a parametric computational study of the Mach 6.0 experiments, the effects of adjacent injectors was found to decrease lateral spreading while increasing the vertical penetration of the injectant plume, and marginally increasing the injectant core decay rate. Matching of the computational results to the experimental results was best achieved when using the Baldwin-Lomax turbulence model without the Degani-Schiff modification. / Ph. D.

Under the assumption of a perfect gas with a constant specific heat ratio, the first approximation of the integral-relations method, which considers the entire shock layer as a single strip, is derived for axisymmetric bodies of arbitrary smooth contour. The resulting differential equations were then applied to a supersonic and hypersonic flow past a paraboloid of revolution. The shock shapes, shock wave detachment distances, locations of sonic lines; and velocity and pressure distributions for the body were calculated for γ = 1.4 and y = 5/3, and at Mach numbers of 3, 4, 6, 10 and 1000. These calculations were carried out on an IBM 1620 electronic computer. The results were compared with those obtained by Van Dyke's inverse method. The agreement between the two methods was found to be good, in view of the fact that only the first approximation of the integral relations method was used. / Master of Science

Liquid and slurry jets were injected through a circular orifice transverse to a M = 3.0 airflow. Mass samples of both jets were taken across the plume 30 injector diameters downstream. Pitot and static pressure surveys were taken across the liquid jet. These data allowed the calculation of distributions across the liquid jet plume of Mach number, air mass flow, liquid-to-air ratio, and momentum flux. A correlation for the liquid concentration in the downstream plane is also presented. In the plume, there is a core region of subsonic airflow carrying two-thirds of the mass collected in the plume. In the core, the liquid mass flow is nearly constant from side-to-side at a given height, and the average velocity of the liquid is only 30 to 60% of the local air velocity. A supersonic mixing region covering two-thirds of the area of the plume surrounds the core region. Comparison with the results from this direct sampling data indicate that correlations developed from photographic techniques are inadequate in determining the jet penetration and width of liquid and slurry jets. The slurry jet showed substantial phase separation. A 30% mass-loaded slurry of 1-5 µm silicon dioxide particles mixed with water was injected, and the local loading varied from a low of 13% at the bottom of the plume to 100% outside the liquid plume. The local loading increased as the jet boundary was approached from any direction. / Master of Science

An experimental study of the ignition of liquid fuels injected transverse to a hot supersonic (M=1.65) air stream was conducted. The liquids considered were kerosene, CS₂ and water as an inert control. The major variables were: air stagnation temperature in the range 1500 to 2300ºF, injectant flow rate and injection angles from 90º to 45º upstream. The experimental observations were: temperature measurements on the wall near the injector and in the flow downstream of injection self-luminosity photographs and infrared photographs taken with a Thermographic camera. Special attention was directed at the behavior of the liquid layer that had previously been found to form near the injector. No unequivocal evidence of ignition of either fuel was found for normal injection at these conditions. However, clear evidence of ignition of CS₂ was found for the upstream injection angle for T_o ≥ 2030°F and 80 ≤ P_j ≤ 135 psi. Higher injection pressures and thus high flow rates failed to produce ignition at any temperature tested. Evidence of CS₂ ignition was found in the infrared photographs and wall and in-stream temperature measurements simultaneously. The infrared photograph indicated possible ignition of the kerosene for upstream injection, but this could not be corroborated with the temperature measurements. / Master of Science

The panel flutter analysis presented herein has been restricted to the problem of an isolated simply supported plate of uniform thickness. The same type of analysis can be applied, however, to other panel configurations. Clamped panels, integrally stiffened panels, arrays of panels, end others should be amenable to treatment by the model approach based on the static aerodynamic approximation. / Master of Science

Tangential supersonic slot injection (M = 1.7) of air into a supersonic air stream (M = 2.91) was studied with a dual hot wire probe. This probe was used to simultaneously determine total temperature and mass flux. Mean profiles across the entire flow field at three axial stations (x/h = 4, 10, 20) were obtained, as well as the temperature and mass flux turbulence intensities. The probe can be used anywhere except at locations where features of the flow are smaller than the separation between the wires (0.18 mm). The calibration of the probe can be repeated with less than a three percent change in calculated Nusselt number. The hot wire probe can also be tuned to obtain a high frequency response. The interaction of a shock wave, caused by a wedge placed on the upper tunnel wall, with the mixing region was investigated. Little change in the mean profiles was observed, but there was a significant increase in turbulence levels. / Master of Science

Large-scale structures were studied in a supersonic stream of air (M = 3) with tangential supersonic slot injection of air (M = 1.7). A dual constant temperature hot-wire probe was used to determine the average structure angles and the characteristic length of the turbulent structures. A zero pressure gradient supersonic boundary layer was studied upstream of the slot injection, and results were compared with previously published data. Structure angles on the order of 50° were obtained throughout the majority of the boundary layer, which was consistent with previously published data. The slot injected flowfield was studied at three axial locations (X/H of 4, 10, 20). Two distinct regions were apparent at each station. A region dominated by the upstream supersonic boundary layer resulted in structure angles on the order of 50°. The mixing region between the slot injected flow (M = 1.7) and the tunnel flow (M = 3) resulted in structure angles on the order of 65°. A compression ramp was used to generate a shock between X/H of 10 and X/H of 20. Structure angles obtained at X/H of 20 appeared unaffected by the streamwise pressure gradient. The characteristic length of the turbulent structures in the supersonic boundary layer and the mixing region of the slot injected flowfield were less than 3.5 mm; however, the characteristic length could not be resolved in this region due to limitations imposed by the frequency response of the hot-wire anemometer systems. / Master of Science

A study was conducted to determine the effect of strut position on the aerodynamic and acoustic performance of a supersonic inlet. The investigated inlet was a prototype 1/14 scale model of a mixed compression, axisymmetric supersonic inlet designed for the high speed civil transport aircraft. A 10.4 cm (4.1 in.) turbofan engine simulator was used in conjunction with the inlet to provide the typical noise signature of a high bypass turbofan engine. Two inlet configurations were investigated in this study. The first configuration was the standard inlet design where the struts are located immediately upstream of the fan. The new configuration has the struts located 3.3 chord length upstream of the fan. The purpose for relocating the strut position was to reduce the flow distortion and radiated noise level. The experiment was conducted at various fan operating conditions in order to simulate aircraft approach. The inlet was tested statically without simulating the inflight speed effects. Steady state measurements were made in order to evaluate the aerodynamic performance of the inlet. The aerodynamic results show that the movement of the struts to a new location allowed the strut wake to diffuse significantly before reaching the fan. This reduced the circumferential distortion parameter by a factor of 2.4, without affecting the pressure recovery of the inlet at a fan Abstract speed of 30,000 rpm (40 PNC). Acoustic measurements were taken in the far field in the 0°-110°sector from the inlet axis. The new configuration of the inlet showed an improved acoustic performance over the standard design. Relocating the struts upstream reduced the blade passing tone by an average of 8 dB (0°-110°) sector, and the overall sound pressure level was lowered by an average of 2.6 dB at a fan speed of 30,000 rpm (40 PNC). / Master of Science