Free-piston Shock Tunnel Research Articles

Quasi-one-dimensional modeling of a free-piston shock tunnel

A Lagrangian formulation for the quasi-one-dimensional modeling of free-piston-driven shock tunnels is described. Three simulations of particular conditions for the T4 shock tunnel are then presented and compared with experimental measurements. The simulations provide very good estimates for both the shock speed and the nozzle-supply pressure obtained after shock reflection and also provide detailed information on the gasdynamic processes over the full length of the facility. This detailed information may be used to identify some of the causes for observed variations in nozzle-supply pressure.

On the compression process in a free-piston shock-tunnel

Preliminary results in the Marseille free-piston shock-tunnel facility are presented. The compression of the driver gas by the piston is studied experimentally for two different geometries of the end of the compression tube. Peak pressures obtained with the end of the compression tube closed, and with bursting of the diaphragm separating the high pressure from the low pressure chamber, are compared with calculated values in the cases of N2 and He as driver gases. A phenomenon of accoustic resonance has been uncovered, generating strong pressure oscillations which, if not properly dealt with, could impair the quality of the useful flow in such a facility.

Experiments on Space Shuttle Orbiter models in a free piston shock tunnel

AbstractHeat transfer and pressure measurements were made on a model of the United States Space Shuttle Orbiter in the University of Queensland's T4 shock tunnel at three angles of attack, with stagnation enthalpies which varied by a factor of 11, from 2·1 MJ/kg to 23 MJ/kg and normal shock Reynolds numbers which varied by a factor of 38, from 2·1 x l04 to 8·l x 105. Leeward pressure results were obtained for comparison with flight data and equilibrium calculations, but the majority of the experiments were conducted to investigate the heat transfer distributions around the shuttle model. Both the windward and leeward heat transfer results exhibit the onset of transition to turbulent flow. The leeward results are compared with flight data as well as with conventional wind tunnel data, and are used to establish trends associated with variation in the major flow and geometry parameters. It was found that although high enthalpy effects could be important, Reynolds number effects played a dominant role in determining the flow.

Measurement of shock-wave/boundary-layer interaction in a free-piston shock tunnel

The short test times available in free-piston shock tunnels necessitate careful examination of experimental techniques. Measurements of shock-wave/bo undary-layer interaction are presented to demonstrate the effec- tiveness of these hypervelocity facilities in studies of a class of flows relevant to scramjet inlets. A two-dimen- sional laminar boundary layer with a freestream Mach number of 5.6, a static pressure of 4 kPa, and a unit Reynolds number of 1.14x 106/m was used as the incoming flow condition on a cold flat plate. A shock wave was generated by a wedge normal to the plate and with a face at 2.84 deg to the freestream direction. It is found that, for this weak interaction, steady flow conditions can be obtained, although the test time is of the order of only 1 ms. HE present high level of interest in aerospace planes and orbital transfer vehicles has focused attention on a num- ber of hypersonic fluid flow problems. One such problem is the unavoidable interaction between shock waves and boundary layers, both in the external flow over a hypersonic vehicle and in the inlet to an air-breathing propulsion system. Shock-wave/boundary-layer interaction is extremely impor- tant in the inlet of the air frame integrated scramjet engine under consideration for the next generation of space launch vehicle. The shock waves that are necessary to compress the inlet flow must glance across the boundary layer developed on the forebody of the vehicle. Similar interactions will occur farther inside the scramjet. The sudden pressure rise through a shock wave causes thickening of the boundary layer in the region of the shock wave and can cause the flow to separate. In closed ducts such as scramjet inlets the extreme result is choking of the flow, with a shock wave standing outside the entrance to the inlet and flow spilling around the sides. Inter- actions can also increase the rate of shear locally in the flow, with hot spots developing on surfaces. These phenomena in scramjet engines are examples of an important class of shock-wave/boundary-layer interactions in which sweepback of the shock wave is a key element. The flow in such interactions is then three dimensional. The configura- tion of interest in this paper is shown in Fig. 1. A planar shock wave generated by a wedge is swept back across a two-dimen- sional boundary layer on a flat plate. The leading edge of the wedge is normal to the plate. If the deflection angle of the wedge is small (less a few degrees), large regions of separated boundary layer generally do not occur and the interaction is regarded as weak. This three-dimensional interaction has been studied extensively13 at freestream Mach numbers of about 3, but investigators disagree over some basic aspects of the flow. There is a dearth of experimental data at hypersonic Mach numbers, particularly at test conditions similar to those expe- rienced in upper atmospheric powered flight and in re-entry.

Stagnation point heat flux in hypersonic high enthalpy flow

The paper describes stagnation point heat flux measurements at a range of enthalpies relevant to re-entry speeds of aero-assisted space transfer vehicles (ASTVs) and proposed space planes, using the Australian National University Free Piston Shock Tunnel T3. The unique feature of these experiments is that they were conducted in the ‘straight through’ (reflectionless) mode which enabled higher enthalpies and densities hitherto unattained.

Free piston shock tunnels: developments and capabilities

The free piston shock tunnel has shown itself to be uniquely capable of adequately simulating real gas effects and has been successfully operating in Australia for more than two decades. As a consequence of the recent interest in hypersonic aerospace planes and general resurgence in hypersonic research, attention has been focussed on this type of facility and free piston shock tunnels have recently been built or being considered in Europe, U.S.A. and elsewhere. It is in the light of these events that this paper discusses the development, performance and capabilities of free piston shock tunnels.

Drag balance for hypervelocity impulse facilities

A new technique is described for measuring drag with 100-/*s rise time on a nonlifting model in a free piston shock tunnel. The technique involves interpretation of the stress waves propagating within the model and its support. A finite element representation and spectral methods are used to obtain a mean square optimal estimate of the time history of the aerodynamic loading. Thus, drag is measured instantaneously and the previous restriction caused by the mechanical time constant of balances is overcome. The effectiveness of the balance is demonstrated by measuring the drag on a cone with 15-deg semivertex angle in nominally Mach 5.6 flow with stagnation enthalpies from 2.6 to 33 MJ/kg. Measurement repeatability of about 10% is achieved.

Measurements of heat transfer to a flat plate in a dissociated high-enthalpy laminar air flow

Heat-transfer rates from a non-equilibrium hypersonic air flow to flat plates at zero and 12° incidence have been measured in a free piston shock tunnel at stagnation enthalpy levels up to 51 MJ kg. Nozzle flow conditions resulted in test section velocities up to 8·1 km 8 and in an experimental regime in which the free stream was chemically frozen and the flat-plate boundary layer was laminar. Estimates of the gas-phase and surface-reaction Damkohler numbers have been made and the heat-transfer results are discussed in this context. At the highest test-section densities non-equilibrium endothermic gas phase reactions involving oxygen atoms in the boundary layer are suggested as a possible mechanism for the observed low heattransfer rates.

The influence of relaxation on shock detachment

By establishing that the length scale governing the detachment of a shock wave from a wedge is the distance from the leading edge to the sonic line, and by considering the view of observers with different length scales, it is predicted that the detachment distance increases gradually with wedge angle for relaxing flow and more rapidly in a perfect gas. Both of these features are confirmed by experiments in the free-piston shock tunnel. The influence of other length scales is discussed. The phenomenon is related to a relaxation effect in which a subsonic layer grows from the translational-rotational shock as the wedge inclination is increased beyond the frozen sonic point.

High-Enthalpy Nonequilibrium Carbon Dioxide Nozzle and Wedge Flows: Experiment and Calculations

A experimental survey is presented of carbon dioxide nozzle expansions from a reservoir at high specific enthalpy, h 0. The aim is to test the chemical model assumed for the gas in numerical calculations and to determine the conditions at the nozzle exit as a function of reservoir condition. Static pressure, which is sensitive to nonequilibrium effects, and pitot pressure are in good agreement with Ref. 1. The gas model is also tested in flows over bodies by comparing optical interferograms of wedge flows with numerical calculations, and it is found that vibrational relaxation, which is unimportant in nozzle flow, plays a significant part here. The results confirm the reliability of the free piston shock tunnel for the study of high-enthalpy flows of carbon dioxide.

High Enthalpy Hypersonic Boundary-Layer Flow

A theoretical and experimental study of an ionizing laminar boundary layer formed by a very high enthalpy flow (in excess of 12 eV per atom or 7000 cal/gm) with allowance for the presence of helium driver gas is described. The theoretical investigation has shown that the use of variable transport properties and their respective derivatives is very important in the solution of equilibrium boundary layer equations of high enthalpy flow. The effect of low level helium contamination on the surface heat transfer rate is minimal. The variation of ionization is much smaller in a chemically frozen boundary layer solution than in an equilibrium boundary layer calculation and consequently, the variation of the transport properties in the case of the former was not essential in the integration. The experiments have been conducted in a free piston shock tunnel, and a detailed study of its nozzle operation, including the effects of low levels of helium driver gas contamination has been made. Neither the extreme solutions of an equilibrium nor of a frozen boundary layer will adequately predict surface heat transfer rate in very high enthalpy flows.

Non-equilibrium dissociating nitrogen flow over a wedge

Experimental results for dissociating nitrogen flow over a wedge, obtained in a free-piston shock tunnel, are described. Interferograms of the flow show clearly the curvature of the shock wave and the rise in fringe shift after the shock associated with the dissociation. It is shown that the shock curvature a t the tip of the wedge can be used to calculate the initial dissociation rate and that it is a more sensitive indication of the rate than can be obtained from fringe shift measurements under the prevailing experimental conditions. Because the free-stream dissociation fraction can be adjusted in the shock tunnel, the dependence on atomic nitrogen concentration of the dissociation rate can be determined by the shock curvature method. A detailed calculation of the flow field by an inverse method, starting from the measured shock shape, shows good agreement with experiments.

Starting Process in the Nozzle of a Nonreflected Shock Tunnel

R experiments have shown that high performance shock tubes will produce flow speeds of 20 km sec. If these shock tubes could be operated as nonreflected shock tunnels, hypersonic flows could be generated at stagnation enthalpies corresponding to speeds exceeding 20 km sec. Test times in such shock tubes are only of the order of 10 //sec, indicating that test time losses occasioned by the nozzle starting process must be kept to a minimum. Furthermore, this must be done without using a diaphragm at the nozzle entrance, as the times available are not sufficient to allow removal of diaphragm fragments from the flow. By using an approximate analysis of the nozzle starting process, together with experiments in a small free piston shock tunnel, it is shown that these conditions are satisfied by establishing steady supersonic flow in the nozzle prior to arrival of the primary shock wave.

Non-equilibrium dissociating nitrogen flow over spheres and circular cylinders

Theoretical results based on the methods of Freeman and Garr & Marrone show that the stand-off distance and flow pattern of non-equilibrium dissociating flow of nitrogen over the front part of a blunt body can be correlated in terms of a single reaction rate parameter ω taking account of parameters describing the speed, density, dissociation and temperature of the free stream. The density pattern, which is sensitive to the reaction rate, consists of two distinct regions dominated by the effects of reaction and pressure respectively. The shape and size of these regions depend on Q. Experimental results obtained by optical interferometry in a free-piston shock tunnel confirm the theoretical results. A scale effect consistent with the induction time phenomenon suggested by Shui, Appleton & Keck modifies the theoretical results considerably in the case of small models.

A study of the free-piston shock tunnel.

A study of the free-piston shock tunnel.

Preliminary results with a free piston shock tunnel

O widely used means of studying high-velocity gas flows is the shock tunnel, which involves a hypersonic nozzle supplied with test gas heated in a shock tube. The speed of the shock in the tube governs the nozzle stagnation enthalpy, and because present techniques limit this speed to about 11,000 fps, maximum enthalpies obtainable correspond to nozzle velocities of about 16,000 fps. The free piston shock tunnel represents an attempt to increase these enthalpies by raising the shock speed. It derives from the free piston shock tube, in which the shock is produced by a driver gas, initially contained in a large tube at relatively low pressure, which is isentropically heated and raised to high pressure by a single stroke of a relatively heavy free piston. The piston speed is sufficiently low that conditions are essentially uniform throughout the gas at any instant during compression. The shock tunnel modification is illustrated in Fig. la and is explained in detail in Ref. 2; only a brief description is given here. The shock tube diaphragm ruptures before the piston compression stroke is complete, at a moment when the piston velocity is such that subsequent movement of the piston face approximately compensates for the flow of gas from the driver volume into the shock tube. Provided that the area ratio of the driver compression tube to the shock tube is large enough to neglect the expansion wave generated in the driver by the bursting diaphragm, the driver conditions then remain approximately constant for a period, which is prolonged by the piston movement. This period is ultimately limited by the need to allow the piston to finally slow down and come to rest before striking the end of the tube. However, with a driver gas of 7 = 1.67, it is theoretically posa)