Hotshot Wind Tunnel Research Articles

Investigation of temperature fields in supersonic flow behind a backward-facing step

The results of numerical modelling and experimental investigations of high-enthalpy turbulent flows in the neighborhood of 90-degree backward-facing steps at the Mach numbers M∞ = 2–4 are presented. The experiments were conducted in the hot-shot wind tunnel IT-302M of ITAM SB RAS. The computations were carried out on the basis of the full Favres-averaged Navier — Stokes equations augmented by the Wilcox turbulence model. The temperature factor influence on the flow structure in the separated zone and temperature distributions was investigated numerically for different Mach numbers. The wall temperature is shown to affect significantly the quantity and sizes of recirculation vortices as well as the temperature distribution in the zone of flow separation and reattachment. The computational results are compared with experimental data on the pressure distribution on the model surface and the wave structure of the flow.

Ignition process evolution at high supersonic velocities in channel

The results of experimental research of multi-injector combustors in the regime of the attached pipe are presented. As a source of high-enthalpy working gas (air), hot shot wind tunnel IT-302M of ITAM, the Siberian Branch of the Russian Academy of Sciences was used. Tests have been carried out at Mach numbers 3, 4 and 5, in a range of change of total temperature from 2000K up to 3000K and static pressure from 0.08MPa up to 0.23MPa. Injector section has been manufactured in two versions with a various relative height of wedge-shaped injectors with parallel fuel injection. Influence of conditions on the entrance of the combustion chamber on ignition and a stable combustion of hydrogen was investigated. Intensive combustion of hydrogen has been received only at Mach numbers 3 and 4. Advantage of injector section with the greater relative height of injectors is revealed. The mechanism of fuel ignition in the combustion chamber of the given configuration was investigated: two-step ignition process including “kindling” and intensive combustion over all channel volume.

Determination of gas-flow composition in hot-shot wind tunnels

A method for gas sampling in hypersonic hot-shot wind tunnels is proposed. A sampler designed on the basis of this method is shown to ensure heat removal for “freezing” the sample and to provide reliable information about the composition of a hypersonic gas flow. A diagnostic system for multispecies gas mixtures on the basis of small-size sensors is developed and manufactured. Testing this system in an IT-302M hypersonic hot-shot wind tunnel shows that the measurements correctly capture the trends in the behavior of oxygen concentration as a function of flow energy, but the values are underestimated approximately by 2% with respect to the true values. Possible reasons for this behavior of the sensor are discussed.

Gas dynamics of a low-pressure chamber in a high-enthalpy short-duration wind tunnel

Gas dynamics of a low-pressure chamber in a double-chamber hotshot wind tunnel is considered. The time needed to fill the second chamber is in good agreement with the value obtained by examining the quasi-steady process of emptying and filling of the high-pressure and low-pressure chambers. The transitional process is strongly affected by wave processes in the channel between the chambers. By changing the place of the throttling grid in this channel, one can change the time of filling of the second chamber (decrease it severalfold). The examined schemes of air input into the second chamber ensure its complete deceleration. The Mach number at the exit of the second chamber (in the windtunnel nozzle throat) is almost constant over the cross section and close to unity.

Structure of Flow over a Hypersonic Inlet with Side Compression Wedges

The e ow over a hypersonic three-dimensional inlet with ramp and side-swept compression wedges was experimentally studied. Two identical inlet models were tested in the blowdown wind tunnel T-313 at Mach numbers M1 =4 and 6 and in the hotshot wind tunnel IT-302M at M1 =6 and 8. Optical visualization of the e ow, near the inlet entrance and in the inlet duct, as well as oil-e lm visualization of the streamlines on compression surfaces were performed. Distributions of static pressure and heat e uxes on these surfaces in typical cross and longitudinal sections, and pitot-pressure distributions at the inlet duct exit, were obtained. Patterns of the supersonic e ow over the compression wedges and in the inlet duct were identie ed. A characteristic feature of these e ows, in particular, is what the compression wedges form shock waves glancing relative to the wedge surfaces. These shocks induce oblique separations of the boundary layer with intense helical vortices propagating far downstream. As a whole, a complex system of multiple shock and expansion waves and vortex structures occurs over the inlet section of external compression, so that the e owe eld at the inlet-duct entrance, in the duct, and at the exit is signie cantly nonuniform. Tests with a natural development of the boundary layer on compression surfaces and with tripping were carried out. “ Grater-type” trips were used, which demonstrated the high efe ciency at high Mach numbers M1 =6‐8. The unit Reynolds numbers in wind tunnels at the same test Mach number M1 o6 were different, Re11 o20£10 6 1/m in T-313 and Re11 =(2‐2.5)£10 6 1/m in IT-302M, respectively, but the boundary-layer tripping ensured identical e ow regimes determined by thestateof the boundary layer. Experimental data obtained in wind tunnels of different operation principles are in good qualitative agreement even with taking into account some quantitative difference mainly attributed to the unit-Reynolds-number difference.

Skin friction gage for measurements in hypersonic flow

A description and results of tests of a new small-scale gage for direct measurement of skin friction force are presented in the paper. The gage design provides separated measurement of longitudinal and transversal component of friction force. Application of this scheme provides high sensitivity and necessary high-frequency response of the gage. The tests of the gage were carried out in a blow down wind tunnel at Mach numbers of 2 and 4 within the range of Reynolds numbers Re x from 0.8 to 5 million and in the hot-shot wind tunnel at Mach number 6 and Reynolds numbers Re x =2.5–10 million. The measurements of skin friction were carried out on a flat plate and on a ramp beyond the shock wave. Simultaneously with the direct measurement of friction in the blow down wind tunnel, the measurements of profiles of average velocities and mass flow rate pulsations were realised. Analysis of measurement errors has shown that the friction gage permits to measure skin friction coefficient on a flat plate with mistake not more than 10%.

Combustion of Liquid and Gaseous Fuels in a Supersonic Combustor

Results of an experimental study of a full scramjet model operating on kerosene, which was performed in an IT-302M hotshot wind tunnel based at the Institute of Theoretical and Applied Mechanics of the Siberian Division of the Russian Academy of Sciences, and an experimental study of a model operating on hydrogen, which was performed in a hotshot wind tunnel with fire heating based at the China Aerodynamic Research and Development Center, are reported. The tests were performed for Mach numbers 5 and 6 for flow parameters close to in-flight conditions. An optimal system for kerosene injection under these conditions was determined, and the thrust characteristics of the engine model were examined. The possibility of controlling kerosene combustion in tests in the short-duration wind tunnel was analyzed, and special features of fuel ignition in a short combustor were considered. Intense combustion of kerosene was achieved with upstream injection of more than 3% of hydrogen, which allowed obtaining effective thrust. The distributions of static pressure and force characteristics of the model in the case of kerosene and hydrogen combustion were compared.

Experimental Study of Scramjet Module

Results of an experimental study of three-dimensional inlets and a scramjetmoduleare presented. The scramjet moduleconsistsofan inletand a combustion chamber.Themoduleisdesigned to study thee owstructure, to obtain the inlet performance, and to study the mutual effects of the combustor and the inlet, fuel ignition and combustion (hydrogenand hydrocarbon fuels ), andtheefe ciency of various methodsof fuelinjection.Thetestswereconducted in a blowdown wind tunnel at Mach numbers from 2 to 6 and unit Reynolds numbers from 8 to 54 £ 106 and in the hot-shot wind tunnel at Mach numbers M1 = 6 and 7.2 and unit Reynolds numbers from 1.4 to 45 £ 10 6 .

High-enthalpy hypersonic project at ONERA

This article deals with the research performed in the framework of the internal ONERA Hyperenthalpic Hypersonic Project. This project involved fifteen researchers specialised in different domains, which are the following: fundamental studies on shock/wave and shock/boundary layers interactions; study of the laminar/turbulent transition; real gas effects and associated physical modelling; the development of non-intrusive optical diagnostic methods usable in cold and hot wind tunnels; flow rebuilding in a hot shot wind tunnel such as F4; conception, realisation and validation of computational fluid dynamics (CFD) solvers.

On the validation of high enthalpy wind tunnel simulations

In the frame of the Ground-to-Flight Transposition methodology for reentry vehicles, Europe has built high enthalpy facilities, the ONERA hot shot wind tunnel F4 and the DLR piston driven shock tube wind tunnel HEG. Both facilities can provide flows at high velocity (5 to 6 km/s) and at high reservoir pressure (several hundreds atmospheres). These facilities are operational since a few years. Despite technological difficulties attached to high thermal loads, they have allowed to study topics of great importance in hypersonics, such as the influence of real gas (i.e. high temperature gas with thermochemistry) on flare-induced boundary layer separation. After discussing on the Ground-to-Flight Transposition methodology in hypersonics, the F4 wind tunnel is presented. Finally, the problem of the validation of high enthalpy wind tunnel experiments is discussed through the analysis of typical experimental results and of the related rebuilding computations.

Velocity Measurement in Hypersonic Flows Using Electron-Beam-Assisted Glow Discharge

A new pseudospark-type electron gun has been utilized in the F4 high-enthalpy, low-density, hot-shot wind tunnel at ONERA. The main goal was to perform accurate measurements of the velocity profile across the boundary layer and the freestream through a time-of-flight principle. In these experiments, an intense pulsed electron beam traces the path of a high-voltage, sustained-glow discharge in some 10 ns. A charge-coupled device camera is opened briefly, 5 μs after the electron gun actuation, to image the position of the luminous column convected by the flow

Flow Characterization in the ONERA F4 High-Enthalpy Wind Tunnel

Experimental results obtained in the ONERA F4 hot-shot wind tunnel have been analyzed to address two important problems for this type of wind tunnel, i.e., the determination of reservoir conditions and the thermochemical nature of the nozzle flow. Numerical tools (one-dimensional and two-dimensional inviscid or viscous, unsteady or space marching codes) are used to reproduce the results of the high-enthalpy experiments. To illustrate this approach, several runs are investigated with different arc-chamber material options, for which diode laser infrared absorption spectrometry (DLAS) has provided freestream velocity, translational temperature, and nitric oxide concentration measurements. The reservoir enthalpy can be determined using spherical and sharp-cone probe heat-transfer-rate measurements with adequate correlations. Recent direct measurements of freestream velocity with electron beam fluorescence time-of-flight technique are used to cross check DLAS and heat-transfer-rate probe results. Concerning the thermochemical nature of F4 nozzle flows, an unexpected conclusion is obtained, as the nozzle wall pressure and translational temperature are observed to be close to equilibrium values at high-enthalpy operating conditions

Condensation of zinc vapor in a supersonic carrier gas

A study of the condensation of a metal vapor in an inert carrier gas is made. Superheated zinc vapor is generated in a hot shot wind tunnel in a helium carrier gas and expanded in a converging-diverging nozzle. Static pressure measurements along the length of the nozzle are made to determine the location of the onset of condensation. A conical nominal Mach 5 (helium) nozzle is employed. The amount of supercooling before the onset of condensation is found to be approximately 430 K, measured along the isentrope over a range of initial zinc mass fractions of .35 to .70 for saturation partial pressures of zinc between 10 psia and 70 psia. The measurements are compared with results of an analysis based on the classical liquid drop model of nucleation. The computed results agree reasonably well with the measurements.

Plasma Sheaths of Models in Hot-Shot Wind Tunnels

The nitrogen flow characteristics around models is studied with a view to obtaining data which determine the physical state of the plasma. Among all possible techniques, the most adaptable to the special hot-shot conditions?electrostatic probes and VHF microwave reflectometry?is chosen. The experimental apparatus and results in nitrogen are described. The ionization is not homogenous during the shot; this result seems to confirm the assumption that the plasma flow is utilizable only after a delay of more than ten milliseconds following the shock formation period. The hot-shot utilization range is examined for this type of experiment.

Computation of electrostatic probe characteristics by quasilinearization.

This paper describes the numerical solution of a mathematical model of a spherical electrostatic probe in a collision-dominated, partially ionized gas and the application of this solution to reduction of probe data gathered from a small hot-shot wind tunnel. In the simple plasma model considered, ionization and recombination are neglected, and current is assumed to flow between the probe and infinity. This mathematical model leads to a sixth-order nonlinear system of ordinary differential equations with two-point boundary conditions. An asymptotic quasineutral solution is used as an initial condition for a solution by quasilinearization. Distribution of charge carrier densities and potentials are presented for several values of dimensionless probe potential and radius, and one case is presented where the ion and electron temperature are not equal. Dimensionless current-voltage curves are presented for a range of dimensionless probe sizes for singly charged ions and equal electron and ion temperatures. A method is developed for analyzing experimental probe data using the numerical solution results.

Errata - "Hotshot wind tunnel force measurement"

Errata - "Hotshot wind tunnel force measurement"

Organisation Et Depouillement Automatique Des Mesures Dans Les Deux Soufflerier a Arc De L'Onera

Several unique conditions are imposed upon measurements made in ONERA hotshot wind tunnels: the duration of flow; the environmental conditions; the circumstances under which the tunnels are used. Flow durations between 40 and 100 milliseconds are produced, and this imposes demands upon transducer performance. Instantaneous power consumed by the tunnels is of the order of 200 megavolt-amperes, and this places the transducers and portions of the readout system in an electrical environment wherein elaborate shielding precautions are required. The contract arrangement under which the tunnels are operated necessitates their use in testing for an optimum number of runs within a half-day of operation, and this necessitates the use of highly automated readout apparatus. This paper describes the solutions to these and other hotshot instrumentation problems. Methods of measurement and readout of pressure, forces, and rates of heat transfer are discussed, and systems for calibration are presented. Many elements of the calibration and readout processes have been made largely automatic. Errors contained among data representing the measurement of forces are sufficiently small as to be considered negligible in the face of minor uncertainties in readout constants. Measurement of pressures is conducted to an overall accuracy (including calibration and readout) of 2.5%. Heat transfer calibrations of 5-10% accuracy are imposed, and overall accuracy of heat transfer data ranges between 15 and 20 percent.

Research and Test Apparatus

The Boeing Airplane Co., Seattle, Washington, U.S.A., recently began regular operation of a new hypervelocity wind tunnel. The electrical are discharge or hotshot wind tunnel is a facility capable of simulating many of the flight conditions which will be encountered by space vehicles. The tunnel is designed to operate in the speed ranges from Mach 10 to 27, at temperatures up to 14,000 deg. F. and can simulate altitude conditions up to 260,000 ft. This addition provides the Boeing Company with aerodynamic research capabilities in all speed regimes from low subsonic to the orbital speed of space vehicles, or from about 20 m.p.h. up to 18,000 m.p.h.