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

A combination of tangential and normal air injection into a Mach 3 airflow was experimentally studied. A rearward facing slot producing tangential injection at a nominal Mach number 1.7 was operated at several different total pressures. An array of transverse tubes of height equal to the slot height and placed just downstream of the slot was operated at two dynamic pressure ratios as well as at Mach 1 and 2.2. Mean flow measurements of static and total pressures were taken up to 20 slot heights downstream from which Mach number, density, velocity and entrainment rates were calculated. Various dimensions and spreading angles of the mixing regions were measured directly from Nanoshadowgraphs and Schlieren photographs. Large eddy structures were produced in several cases, leading to increased entrainment of the free stream. For some cases, heated air was injected through the normal tubes, and the jet total temperature decay was measured downstream. It can be seen from the data as a whole that the mixing rate can be significantly increased by the combined tangential-normal injection design over tangential slot injection alone, with up to as much as 92% more entrained mass. / Master of Science

Experimental results of hot-·wire measurements of turbulence in a supersonic shear layer are presented. A multiple-overheat single-wire method based on high-speed, computer-controlled sampling was used in order to obtain large sample sets in short time intervals. The Mach 3 main stream had a total pressure of 95 psia and a Re/cm of 2 x 10⁶. The Mach 1.7 secondary stream was injected parallel to the main flow at 10.7 psia, resulting in a slightly overexpanded flow. The injection slot had a height of 1.2 cm. The flow was monitored at four streamwise stations (x/H = .25, 4, 10, 20). The total temperature of the injected stream was varied from 300 K (equal to that of the main stream) to 420 K. The effects on the flow of these changes in density, velocity and temperature was investigated. Measured quantities included the mean and rms levels of the mass flux and total temperature. Nanosecond Shadowgraphs and spark Schlieren photographs were taken. The Favre-averaged velocity fluctuation, u^H, was also calculated. Small increases in absolute turbulence levels were seen in the heated flow, and the location of the maximum turbulence shifted significantly. Heating greatly enhanced the mixing in the shear layer and in the injected flow by the last station. An investigation of shock interaction with the shear layer generally resulted in elevated turbulence, but had little effect on. The measurement technique and reduction method proved accurate in the unheated flow, but greater uncertainty was found for the heated injection case. / Master of Science

This theoretical and experimental investigation has proven that accurate results can be obtained by using the hydraulic analogy. The analogy can be used in either the subsonic, transonic or supersonic range. The theory was intended to be general and showed that an exact analogy exists between free surface water flow and two dimensional gas flow. By theoretical investigation, a new channel was proposed and experiments were conducted. Remarkable correlation between theory and experiment was found and it is believed that the method presented in this thesis represents a closer approximation to the actual flow of air than any of the other present day analogies. It has been shown that by using the analogy, the pressure, density, and temperature can be easily obtained. In addition, the shock wave phenomenon can be easily observed. This analogy would be an excellent took in demonstrating the different types of flow around bodies. Analogy research in the field of supersonic aerodynamics is not only justified, but also necessary since the present day experimental technique in wind tunnels present economical and scientific difficulties. A large amount of knowledge can be obtained from analogy research in this field. The analogy described in the paper gives accurate results; the proposed apparatus is inexpensive, and experiments can be performed in a very short time. / Master of Science

Master of Science

A finite difference method of investigating the shock layer in the vicinity of the stagnation point in a two-dimensional supersonic flow past a circular cylinder is analyzed. This method differs from existing methods which have been used in the past for solving this type of problem in that it uses the properties along the stagnation streamline as initial conditions. However, the method can still be classified as one of the direct type. A potential flow solution along the stagnation streamline is used to approximate the initial conditions. The results are compared with existing solutions and available experimental data. / Master of Science

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

Steady, inviscid, supersonic flow past conical wings is studied within the context of irrotational, nonlinear theory. An efficient numerical method is developed to calculate cones of arbitrary section at incidence. The method is fully conservative and implements a body conforming mesh generator. The conical potential is assumed to have its best linear variation inside each cell; a secondary interlocking cell system is used to establish the flux balance required to conserve mass. In regions of supersonic cross flow, the discretization scheme is desymmetrized by adding the appropriate artificial viscosity in conservation form. The algorithm is nearly an order of magnitude faster than present Euler methods. It predicts known results as long as the flow Mach numbers normal to the shock waves are near 1; qualitative features, such as nodal point lift-off, are also predicted correctly. Results for circular and thin elliptic cones are shown to compare very well with calculations using Euler equations. This algorithm is then implemented in the design of conical wings to be free from shock waves terminating embedded supersonic zones adjacent to the body. This is accomplished by generating a smooth cross-flow sonic surface by using a fictitious gas law that makes the governing equation elliptic inside the cross-flow sonic surface. The shape of the wing required to provide this shock-free flow, if such a flow is consistent with the sonic surface data, is found by solving the Cauchy problem inside the sonic surface using the data on this surface and, of course, the correct gas law. This design procedure is then demonstrated using the simple case of a circular cone at angle of attack.