Showing papers on "Supersonic speed published in 1969"

Features of plasma transport in the upper atmosphere.

Abstract: Steady, ambipolar, low and high speed plasma transport equations developed for terrestrial ionosphere, discussing boundary conditions for supersonic and subsonic flows

The interaction of shock waves and heat addition in the design of supersonic combustors

Abstract: Analytical models are needed that will permit the prediction of thrust efficiency of a supersonic combustor as a function of combustor geometry and heat release. An earlier, more general procedure for the one-dimensional analysis of supersonic combustors was developed on the basis of exponential pressure-area dependence. This paper extends that analysis by establishing three regimes for specific pseudo-one-dimensional solutions for thrust efficiency: (1) a low-heat-release regime in which the flow can be considered essentially shock-free; (2) an intermediate regime in which an oblique shock wave is sustained with a pressure ratio equal to or greater than that required for turbulent boundary-layer separation; and (3) a high-heat-release regime in which the combustion is preceded by a normal shock. The key to solution is the formulation for the wall-pressure force needed for the momentum equation. With it, a set of integral equations can be solved simultaneously for relationships between pressure ratio, total temperature ratio and area ratio across the combustor in the three regimes. It is argued that the most probable solutions are those that yield the lowest back pressure without violating the entropy limit (second law of thermodynamics). It is shown that such limiting processes, including shock waves, can yield higher over-all efficiencies than shock-free solutions, because the total pressure loss across the shock is more than compensated by the reduction in total pressure loss due to heat addition, which occurs at a lower Mach number. Results of a typical substantiating experiment falling in this class are presented, for a combustor-inlet Mach number of 3.2 and an alkyl borane fuel.

Theory of Laminar Viscous-Inviscid Interactions in Supersonic Flow

Abstract: This investigation is concerned with those fluid-mechanical problems in which the pressure distribution is determined by the interaction between an external, supersonic inviscid flow and an inner, laminar viscous layer. The boundary-layer approximations are assumed to remain valid throughout the viscous region, and the integral or moment method of Lees and Reeves, extended to include flows with heat transfer; is used in the analysis. The general features of interacting flows are established, including the important distinctions between subcritical and supercritical viscous layers. The eigensolution representing self-induced boundary-layer flow along a semi-infinite flat plate is determined, and a consistent set of departure conditions is derived for determining solutions to interactions caused by external disturbances. Complete viscous-inviscid interactions are discussed in detail, with emphasis on methods of solution for both subcritical and supercritical flows. The method is also shown to be capable of predicting the laminar flow field in the near wake of blunt bodies. Results of the present theory are shown to be in good agreement with the measurements of Lewis for boundary-layer separation in adiabatic and non-adiabatic compression corners, and with the near-wake experiments of Dewey and McCarthy for adiabatic flow over a circular cylinder. Extensions of the method to flows with mass injection at the surface and to subsonic interactions are indicated.

Continuous Supersonic Plasma Wind Tunnel

Abstract: The normal magnetic field configuration of a Q device has been modified to obtain a “magnetic Laval nozzle.” Continuous supersonic plasma “winds” are obtained with March numbers ∼3. The magnetic nozzle appears well suited for the study of the interaction of supersonic plasma “winds” with either material or magnetic obstacles.

Acoustic energy flux from shock-turbulence interaction.

Abstract: The analysis of the sound field generated by the passage of isotropic turbulence through a shock of finite strength (Ribner 1953, 1954) has been extended to provide the flux of acoustic energy emanating from unit area on the downstream side of the shock. This is motivated by the problem of estimating the sound power emerging from a supersonic jet containing shock waves. The energy flux is found to vary almost linearly with shock density ratio, reaching a maximum at infinite Mach number of 0[sdot ]062 of the flux of turbulence kinetic energy convected into unit area of the shock.Direct comparison with a result obtained by Lighthill (1953) is misleading. His energy relations, reckoned relative to a frame moving with the fluid, must be converted to the shock-fixed frame used herein. The converted results of his theory (weak shocks) and the results of our theory (arbitrary shocks) appear to show a similar asymptotic behaviour for vanishing shock strength; they diverge with increasing shock strength.

Crack propagation at supersonic velocities

Abstract: A theoretical model for laser-induced supersonic cracks in crystals with weak cleavage planes is developed. The crack is pictured as being driven by a laser-induced expanding plasma. The theory predicts the crack length-time history to be expected for a given energy-time input curve. The conditions for laser-induced supersonic crack production are seen to be a large input energy relative to fracture energy, a short loading time, and a weak fracture plane. The theoretical calculations are in rough agreement with experimental observations.

A supersonic mixing chemical laser

Abstract: The reaction F + HCl → HF† + Cl has been used to produce cw chemical laser action. This has been accomplished by mixing two supersonic streams of the reactant species in a shock tunnel. The time duration of the laser action is two orders of magnitude longer than the natural laser pulse in a non‐flowing system, i.e., it is limited only by the flow time of the shock tunnel.

Theoretical consideration of combustion effects on base pressure in supersonic flight

Abstract: A theoretical treatment is given for the behavior of base pressure on a two-dimensional body in supersonic flight when the inviscid stream adjacent to the turbulent reattachment zone is subjected to combustion effects. The analysis, which rests upon a simplified treatment of the combustion zone and the Crocco-Lees mixing theory for the reattachment region is intended to give an upper-limit estimate of thrust levels attainable on the base. Previous experimental work in this area is reviewed and is shown to suffer in performance because of experiment design; the maximum performance levels could not be expected and further work in this area is indicated. Analytically, it is shown that, despite the apperance of many base flow and combustion parameters, the specific impulse is primarily a function of only the air-fuel ratio, combustion length-body base height ratio, and the free stream Mach number. The major influence of combustion on the reattaching flow is shown to be the introduction of a new characteristics length, the combustion length, and the action of this scale in constraining the reattachment length.

A law of the wall for compressible turbulent boundary layers with air injection

Abstract: A considerable body of experimental data now exists concerning turbulent boundary layers with air injection at the wall, both at subsonic and at supersonic speeds. In the present report these data for Mach numbers up to 6·5 have been analyzed to find the parameters which occur in the law of the wall as deduced from mixing-length theory. Although the absolute values of the parameters are subject to error because of the lack of accurate skin-friction measurements, the trends of these parameters with Mach number and injection mass flow are clearly defined.

Theoretical and experimental study of supersonic steady flow around inclined bodies of revolution

Abstract: Three dimensional hypersonic steady flow around blunt and pointed cones at nonzero angles of attack calculated by method of characteristics

Jet Interaction Control Effectiveness for Subsonic and Supersonic Flight

Abstract: : Interference effects between a highly underexpanded, sonic or supersonic jet in a subsonic or supersonic crossflow, and the surface from which the jet exhausts are examined. For subsonic freestream Mach numbers, existing data are examined and correlated. Various semi-empirical models to represent the interference pressure distribution on flat plates are then developed. For supersonic freestream Mach numbers, a computer program for calculating jet interference effects on axisymmetric bodies at angle of attack is described. Interference effects between the jet plume and control fins on a cruciform missile are analyzed. A semi-empirical model of the jet in a crossflow, valid at large distances from the nozzle is developed. The results of this model are then used to compute interference forces and moments on fins located aft of the nozzle, both for subsonic and for supersonic freestream Mach numbers.

Development of an Advanced Annular Combustor

Abstract: The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

Application of Godunov's Method to Blunt-Body Calculations

Abstract: A modification of Godunov's finite-difference method for computation of steady or unsteady flowfields about planar or axisymmetric blunt bodies is presented. Results for ideal gas flows over two body shapes are shown. The first case is that of a sphere in a Mach number 10.0 freestream, and is used to evaluate the accuracy of predicted results by comparison with other computations. Pressures and stagnation point velocity gradient are found to be accurate to within several percent and are improved in the stagnation region over the values previously achieved with the method. The second test shape is that of a spherically blunted cone both with sharp and rounded shoulders. Two freestream conditions were employed, one resulting in an essentially subsonic shock layer and the other in an essentially supersonic shock layer. Body pressures and shock shapes are presented and compared with available data.

Boundary-Layer Tripping with Emphasis on Hypersonic Flows

Abstract: Experiments on the effect of trip geometry, size, and location on the position of transition at local Mach numbers up to 8.5 are presented. The pressure drag of the trip is investigated at local Mach numbers of 4.7 and 5.5. Based on test results, a flow model was constructed which includes trip-produced multiple vortex filaments similar to those found at supersonic speeds that are assumed to be responsible for introducing the disturbances that lead to transition. The Reynolds number based on the distance from the leading edge to the roughness position as well as Mach number and wall-to-total temperature ratio should be considered in choosing the smallest effective trip size. The effect of trip shape on the position of transition at hypersonic speeds is small; however, certain shapes exhibit the advantageous characteristic of having drag coefficients which were relatively independent of roughness height over a restricted range of the variables tested.

Experiments on shock stand-off distance in non-equilibrium flow

Abstract: An experimental investigation of the non-equilibrium behaviour of the shock stand-off distance ahead of spheres at low supersonic Mach number is reported. An intermittent wind tunnel operating with a reacting gas mixture with one non-equilibrium mode is described. A non-equilibrium parameter after Damkohler is determined for the flow and the experiments cover a range of this variable from near-frozen to near-equilibrium states. The shock stand-off distance is measured and found to vary as expected between these two bounds with a low value at equilibrium. With all other variables governing shock standoff distance held constant, such measurements can also be used to determine the relaxation time of the non-equilibrium mode.

Method and structure for reduction of shock waves

Abstract: A method of and apparatus for reducing shock waves created by solid bodies having supersonic speed relative to a gas, and reducing and even eliminating the sonic boom which accompanies the shock waves. The nose or leading portion of the moving body, be it a body of revolution, or an airfoil lifting body, or be it another moving body, is provided with blunt configuration at its leading portion to create a detached ''''normal'''' shock wave. A three-step thermodynamic cycle is applied to the gas (i.e., air) flow surrounding the moving body. The three steps are (1) compression of the gas, caused by the moving body, (2) heating of the compressed gas by addition of heat, and (3) expansion of the compressed gas to hear its original pressure. In air, for example, a blunt nose flying body at supersonic speeds creates a high compression zone (stagnation zone) between the shock wave and the nose of the body. Heat is applied, e.g., by burning fuel, electricity, or nuclear radiation, to the stagnation zone of high pressure, resulting in a substantial decrease of static pressure immediately in front of the blunt nose. The heat can be applied to the compressed air zone from forwardly projecting structures suitably shaped to create only low intensity shock waves. The applied heat decreases the static pressure at the nose zone, in turn decreasing the high air drag, and also decreasing the intensity of the shock wave ahead of the flying body, thereby decreasing the intensity of the part of the shock wave which reaches the ground, resulting in a reduction in sonic boom.

Acoustic measurements in supersonic transitional boundary layers

Abstract: : Surface pressure fluctuations associated with transitional and turbulent boundary-layer flows on a sharp, slender cone at supersonic Mach numbers were experimentally investigated in a 40- by 40-in. supersonic wind tunnel using a flush-mounted 0.25-in.-diam microphone. The results at Mach numbers 3 and 4 demonstrate the feasibility of locating microphones onboard wind tunnel test models to measure overall pressure fluctuations and power spectral distributions in transitional and fully developed turbulent flows. Transition Reynolds numbers determined using a surface microphone are compared with two other established methods of detection. Selected boundary-layer pressure fluctuation characteristics (power spectral density and root-mean-square values) and transition profiles are presented. Methods of data acquisition and analysis are discussed.

Analysis of fluid dynamics of supersonic combustion process controlled by mixing

Abstract: The fluid dynamics of supersonic combustion is discussed. The interference between the combustion process and the supersonic flow secondary to the combustion region is described. Multiple injector flow fields are described from the point of view of mixing and of interaction with the external flow. It is shown that the selection of injector location and combustion process can be utilized to produce compression waves of controlled strength. Such waves can be utilized to reduce the flow Mach number in front of subsequent injectors. This effect is called thermal compression. Engineering criteria for utilization of thermal compression are presented An example of such utilization is described. The example indicates that the interaction between combustion and geometry is of primary importance for the fluid-dynamic process. The effect of this interaction cannot be accounted for by a simple one-dimensional analysis. Only a judicious combination of mixing analyses and more complex analyses, that takes into account the formation and propagation of the waves due to combustion, can give detailed qualitative information on the fluid dynamics of supersonic combustion.

Turbulent Mean-Flow Measurements in a Two-Dimensional Supersonic Wake

Abstract: Mean‐flow measurements have been made in the supersonic wake of a very slender, two‐dimensional body at zero incidence and heat transfer rate. Excellent correlation of the lateral distribution of axial velocity is obtained when the lateral distance is divided by the transverse wake scale which is formed from the measured velocity defect. The latter is found to decay as the inverse square root of distance for the entire range of about 1850 virtual model thicknesses mapped; thus, the transverse scale increases as the square root of distance. In the latter half of this range the lateral distribution of static temperature also appears to correlate in the same lateral coordinate, and the temperature defect also decays as the inverse square root of distance. It is demonstrated that these results are accurately predictable from a basic similarity analysis beginning with Townsend's measured velocity decay on the axis; the turbulent Reynolds number of 13.0 agrees closely with Townsend's 12.5. The corresponding Prandtl number found lies in the range from 0.65 to 0.70.

Aerodynamically compensated static pressure tube

Abstract: A strut-mounted static pressure tube which is aerodynamically compensated to obtain desired static pressure measurements under supersonic flight conditions as well as under subsonic flight conditions. Compensation of the measured static pressure at the probe is achieved for both subsonic and supersonic operation of aircraft with one instrument.