Free-piston Shock Tunnel Research Articles

PLIF thermometry in shock tunnel flows using a Raman-shifted tunable excimer laser

Planar laser-induced fluorescence is performed in a free-piston shock tunnel by using a Raman-shifted tunable excimer laser to excite nitric oxide molecules in the flow. Two different flowfields are examined to test the difficulties associated with applying the technique to shock tunnels: the bluff body flow produced by a 25 mm diameter cylinder; and the oblique shock and expansion fan produced by a 35° half-angle wedge. For the cylinder, the maximum flow enthalpy was limited to 4.1 MJ kg\(^{-1}\) due to high flow luminosity which is produced by metallic contaminants in the flow. A reflective filter is used to reduce the influence of flow luminosity making these measurements feasible. Freestream temperature measurements are in excellent agreement with those predicted from numerical flow calculations. Large uncertainties were observed for the high-temperature post-shock results. Several higher enthalpy shots (14 MJ kg\(^{-1}\)) were also performed with the wedge and showed an insignificant amount of contaminant emission.

Improvement of a free piston driver for a high-enthalpy shock tunnel

In order to improve the operation of a high-enthalpy free piston shock tunnel its tuned operation was studied analytically and experimentally. First, the piston motion in the free piston driver tube was analytically solved by proposing a simple piston/gasdynamic model, and the tuned operation condition was formulated as an eigenvalue with which the piston has sufficiently high speed at the moment of diaphragm rupture, so as to maintain a constant driver gas pressure, and reduces its speed to come to rest when very closely approaching the end of the driver tube. Second, the result of this analysis was validated by its comparison with experiments which were conducted in the medium-sized free piston shock tunnel HEK installed at the NAL Kakuda Research Center. By observing the detail of piston landing at the end of the driver tube the present tuned operation was found to be successfully achieved with the operating condition given here. Its advantages in improving the pressure recovery factor and in enhancing the stagnation enthalpy were successfully demonstrated.

Fundamental Studies on Heat Load of High Enthalpy Shock Tunnel. 1st Report. Numerical Analysis Method and Its Prediction Accuracy for a Flow Field.

It is very important in high enthalpy shock tunnels to predict heating conditions of the shock tube end where the tube wall is exposed to high temperature. The final goals of these studies are to establish methods to predict thermal conditions to which a tube end is subjected, and to develop thermal protection which will be suitable for high enthalpy shock tunnels. In the present report, a numerical flow analysis method combined with a heat conduction analysis of tube wall was developed to calculate the flow and wall terperature conditions in a shock tube. A numerical procedure was proposed which includes calculation of the mixing region of driver and test gases, and a friction factor model is presented which considers the change of visconsity under high terperatures. Experiments were carried out using the free-piston shock tunnel T5 of the California Institute of Technology. Comparison of calculated and experimental results of pressure histories in the shock tube showed that the present computer program has enough accuracy to calculate the flow field at the shock tube end.

Establishment of steady separated flow over a compression-corner in a free-piston shock tunnel

The time required to establish steady separated compression–corner flow is examined under hypervelocity conditions in a free-piston shock tunnel. This time is reasonably well described using previous perfect gas analyses. The results suggest that, provided the nozzle reservoir enthalpy is 20 MJ kg\(^{-1}\) or less, there is sufficient time to establish steady separated flow before driver gas contamination becomes a significant problem in the present facility.

The interaction of a shock wave with a laminar boundary layer at a compression corner in high-enthalpy flows including real gas effects

The high-enthalpy, hypersonic flow over a compression corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg−1 to 19 MJ kg−1, with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7×104 to 2.7×105. The flow was laminar throughout. The experimental data compared well with theories valid for perfect gas flow and with other relevant low-to-moderate enthalpy data, suggesting that for the current experimental conditions, the real gas effects on shock wave/boundary layer interaction are negligible. The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the compression corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction. The increase in peak heat transfer is attributed to the reduction in boundary layer displacement thickness due to real gas effects.

Fuel Stagnation Temperature Effects on Mixing with Supersonic Combustion Flows

An experimental study of the effect of fuel stagnation temperature on mixing in a supersonic hydrogen-air flame is described, The combustor consisted of a constant-area rectangular duct with a centrally located fuel-injection strut that spanned the width. A high-enthalpy stream of air was supplied by a free-piston shock tunnel, and heated hydrogen fuel, supplied by a gun-tunnel, was injected into the freestream as a coflowing planar jet. The freestream total enthalpies were 5.6, 6.5, and 9 MJ/kg, and fuel stagnation temperatures were 300, 450, and 700 K, Raising the fuel stagnation temperature increased the fuel velocity to be near that of the airstream and resulted in a decrease in the mixing rate, Even as the fuel and air velocities became equal, significant mixing still occurred because of a large difference in density, Increasing the freestream enthalpy reduced the difference between the initial air temperature and the adiabatic flame temperature, which in turn reduced the heat addition, and subsequently, the amount of pressure rise in the duct.

Turbulent spots in boundary layers in a free-piston shock-tunnel flow

In this paper, experiments to detect turbulent spots in the transitional boundary layers, formed on a flat plate in a free-piston shock tunnel flow, are reported. Experiments indicate that thin-film heat-transfer gauges are suitable for identifying turbulent-spot activity and can be used to identify parameters such as the convection rate of spots and the intermittency of turbulence.

96/06405 Species measurements in a hypersonic hydrogen-air, combustion wake

A continuously sampling, time-of-flight mass spectrometer has been used to measure relative species concentrations in a two-dimensional, hydrogen-air combustion wake at mainstream Mach numbers exceeding 5. The experiments, which were conducted in a free piston shock tunnel, yielded distributions of hydrogen, oxygen, nitrogen, water and nitric oxide at stagnation enthalpies ranging from 5.6 MJ kg(exp -1) to 1.2 MJ kg(exp -1) and at a distance of approximately 100 times the thickness of the initial hydrogen jet. The amount of hydrogen that was mixed in stoichiometric proportions was approximately independent of the stagnation enthalpy, in spite of the fact that the proportion of hydrogen in the wake increased with stagnation enthalpy. Roughly 50 percent of the mixed hydrogen underwent combustion at the highest enthalpy. The proportion of hydrogen reacting to water could be approximately predicted using reaction rates based on mainstream temperatures.

Species measurements in a hypersonic, hydrogen-air, combustion wake

A continuously sampling, time-of-flight mass spectrometer has been used to measure relative species concentrations in a two-dimensional, hydrogen-air combustion wake at mainstream Mach numbers exceeding 5. The experiments, in a free piston shock tunnel, yielded distributions of hydrogen, oxygen, nitrogen, water, and nitric oxide at stagnation enthalpies ranging from 5.6 MJ kg −1 to 12.2 MJ kg −1 and at a distance of approximately 100 times the thickness of the initial hydrogen jet. The amount of hydrogen mixed in stoichiometric proportions was approximately independent of the stagnation enthalpy, despite the fact that the proportion of hydrogen in the wake was increased with stagnation enthalpy. Roughly 50% of the mixed hydrogen underwent combustion at the highest enthalpy. The proportion of hydrogen reacting to water could be approximately predicted using reaction rates based on mainstream temperatures.

Upstream influence and peak heating in hypervelocity shock wave/boundary-layer interaction

Two important characteristics of separated flow, namely, the scale of interaction and the peak heat transfer, have been examined under high-enthalpy, hypersonic conditions. Experimental data have been obtained in a free-piston shock tunnel using a compression corner model. The data are examined in terms of established perfect gas theories, and are seen to be in reasonable agreement with these theories and also with data from perfect gas experiments. This indicates that, for the present conditions, the real gas effects on separation and reattachment are small.

High-enthalpy, hypersonic compression corner flow

The results of an experimental investigation of high-enthalpy, hypersonic flow over sharp leading-edge compression corners are presented and discussed. In particular, the possible effects of real gas behavior are examined. Measurements have been made of the heat transfer and pressure distributions for flat plate and compression corner flow. Some flow visualization data have also been obtained. Test flows were generated using a free-piston shock tunnel operating in the reflected mode. The reservoir enthalpy ranged from 3 to 19 MJ kg -1 , giving freestream speeds of 2.3-5.5 km s -1 . For these conditions, the flow remains laminar throughout. The flat plate data for both high- and low-enthalpy flows are in agreement with the reference enthalpy method for heat transfer and the weak interaction theory for pressure. Also, the measured flat plate boundary-layer thickness compares well with an expression strictly valid for perfect gas flows only. The high- and low-enthalpy compression corner flows have upstream influence and plateau pressure behavior similar to perfect gas flow. That is, real gas effects for the present flows appear to be negligible. This is consistent with the essentially chemically frozen viscous and inviscid flow upstream of the interaction.

An experimental investigation of hypervelocity flow in a conical nozzle

The flow in a conical nozzle is examined experimentally for a range of hypervelocity conditions in a free-piston shock tunnel. The pitot pressure levels compare reasonably well with an inviscid numerical prediction which includes a correction for the growth of the nozzle wall boundary layer. The size of the nozzle wall boundary layer seems to be well predicted by semi-empirical expressions developed for perfect gas flows, as do data from other free-piston shock tunnels.

Three-component force balance for flows of millisecond duration

The authors investigate the feasibility of a new type of multiple-component force balance for measurements on models in hypervelocity hows of millisecond duration. The balance extends the concept of the single component stress-wave force balance to measurement of axial force, normal force, and pitching moment. Numerical modeling of the performance of the balance shows that coupled deconvolution techniques can be used to decouple the signals from axial strain measurements in the balance to determine the applied loads. Experiments performed in the T4 free-piston shock tunnel on a sharp cone at incidence indicate that the prototype balance performs satisfactorily, forces being measured to within 11% of their theoretical values, and the location of the line of force being measured to within 2.1% of the theoretical location as a fraction of model chord.

Transverse jet mixing and combustion experiments in hypervelocity flows

Transverse jet mixing and combustion experiments at a freestream velocity of 5 km/s have been conducted in the T5 free-piston shock tunnel at GALCIT. The experiments were performed in a 50 by 25 mm combustor model using a flush-mounted, Mach 1.7 single injector introducing the fuel at an angle of 15 deg relative to the main flow. The test conditions cover a range of freestream pressures from low values chosen to match previous experiments with the same duct in the HYPULSE expansion tube facility at GASL, to high pressures corresponding to the conditions closer to those in a real scramjet combustor. Tests with cold and hot hydrogen fuel were performed, the latter to better reproduce the fuel conditions of a scramjet propulsion system where the hydrogen will have to be used first as a cooling fluid. A small combustion-driven shock tunnel was used simultaneously with T5 to supply the hot fuel.

Experiments on supersonic combustion ramjet propulsion in a shock tunnel

Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg−1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg−1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pilot and static pressure measurements showed that the combustion was supersonic.The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.

Measurement techniques in high-enthalpy hypersonic facilities

Interest in hypersonic aerodynamic braking and aerospace planes has led to the development of new high-enthalpy hypersonic test facilities. This paper describes available techniques and future developments for making measurements in these facilities. Emphasis is on instrumentation for use with short-duration flows in free-piston shock tunnels. Issues of free-stream calibration, surface measurements on models, flow visualization, laser diagnostics, and accuracy are addressed, with references made to representative experimental studies. The paper concludes with a summary of future research needs.

204 低速VISARを用いた圧縮管管内のピストンの速度履歴と計測に関する研究

A free piston shock tunnel (FPST) is one of the most useful test facilities for generating high enthalpy flow. Therefore, the FPST has been intensively used in conjunction with the development of re-entry vehicles and SCRAM jet engine. In this study, piston motion in a compression tube of a free piston compression tube was measured continuously in order to obtain the quantitative piston speed of a FPST by means of a high resolution interferometer and results were compared the computation based on concept of Tuned Operation Conditionwhich uses a quasi-1D ENO scheme with 4th order accuracyGood agreement was obtained between measured numerical results and predictions. A double exposure holographic interferometric visualization of a shock wave released from the compression tube was carried out.

The upgraded ballistic range facilities at AEDC

A major upgrade of the Ballistic Range/Track G at AEDC has recently been completed providing a hypervelocity testing complex unequaled in the world. The new AEDC Hypervelocity Testing Complex consists of:1. A large 84-mm (3.3-in.)-bore two-stage light-gas launcher, capable of launching models into the 930-ft-long Range G, either free-flight or with track guidance and model recovery.2. A 63.5-mm (2.5-in.) two-stage light-gas launcher, with similar performance to the previous 13 launcher, capable of launching models into a dedicated Impact Range.3. A high-performance free-piston shock tunnel which utilizes the 2.5-in. Impact Launcher as the driver. The two hypervelocity ranges (1 and 2) are described and examples of different testing diagnostics are given. These diagnostics cover the spectrum from launcher to a variety of impact events.

Numerical study of free-piston shock tunnel performance

A free-piston shock tunnel (FPST) is one of the most useful ground testing facilities for hypervelocity flow research of re-entry vehicles and scramjet engines. For an efficient operation with tuned piston motion, the design of facility and the comprehension of the physical phenomena in a FPST, a numerical simulation which can properly predicts the flow with actual losses is required. But there are few successful numerical methods which can simulate its overall performance. In the present study, numerical method was developed by using the KRC shock capturing scheme and by modeling the flow losses in suitable forms for a quasi-1D numerical computation. The present numerical results were compared with the data obtained in two different facilities, T4 and T5. The applicability of the present numerical method as a design tool is discussed briefly.

Transition of compressible high enthalpy boundary layer flow over a flat plate

AbstractThis paper presents the results of an experimental investigation into the characteristics of boundary layer transition to turbulence in hypervelocity air flows. A series of experiments was conducted using a flat plate model, equipped with static pressure and thin film heat transfer transducers, in a free piston shock tunnel. Transition was observed in the stagnation enthalpy range of 2·35 to 19·2 MJ/kg. The transition Reynolds number correlates well with the unit Reynolds number through a simple empirical relation. The influences of Mach number, pressure and wall cooling are examined. The measured heat transfer rates in laminar and turbulent regions are compared with empirical predictions. Freestream disturbances of the test flow were also measured and analysed.