Free-piston Shock Tunnel Research Articles

Aerodynamic characteristics of a free-flight scramjet vehicle in shock tunnel

A multi-component aerodynamic test for an airframe-engine integrated scramjet vehicle model was conducted in the free-piston shock tunnel HIEST. A free-flight force measurement technique was applied to the scramjet vehicle model named MoDKI. A new method using multiple piezoelectric accelerometers was developed based on overdetermined system analysis. Its unique features are the following: (1) The accelerometer’s mounting location can be more flexible. (2) The measurement precision is predicted to be improved by increasing the number of accelerometers. (3) The angular acceleration can be obtained with single-axis translational accelerometers instead of gyroscopes. (4) Through the averaging process of the multiple accelerometers, model natural vibration is expected to be mitigated. With eight model-onboard single-axis accelerometers, the three-component aerodynamic coefficients (Drag, Lift, and Pitching moment) of MoDKI were successfully measured at the angle of attack from 0.7 to 3.4 degrees under a Mach 8 free-stream test flow condition. A linear regression fitting revealed a 95% prediction interval as the measurement precision of each aerodynamic coefficient.Graphical abstract

Influence of shock attenuation on tailored operation in free piston shock tunnels

The free piston shock tunnel is a type of shock tunnel with high performance. For this type of tunnel, the influence mechanism of shock wave attenuation on tailored operation is explored by numerical simulation and theoretical analysis. By introducing the normalized velocity, the simple constraint equation for shock wave under the tailored operation is deduced. Moreover, the real gas effect is also taken into account in this equation. Based on the equation, the tailored operation of shock tunnels can be predicted with very few calculations. The present study shows that the change rate of the thermodynamic state of the gas behind the shock wave is inconsistent with the attenuation rate of the shock wave, which is the fundamental reason why the wind tunnel achieves tailored operation at a lower Mach number of shock waves. This lower Mach number of shock waves differs from the corresponding ideal value by a factor, which is about the square root of shock attenuation rate.

Kinetic modeling of unsteady hypersonic flows over a tick geometry

Hypersonic separated flows over the so-called “tick” geometry have been studied using the time-accurate direct simulation Monte Carlo (DSMC) method and global linear theory. The free stream condition for two experimental cases studied in the free-piston shock tunnel (named T-ADFA) was modeled. These two cases span a Knudsen number from transitional to continuum, a Mach number of about 10, a free stream enthalpy from 10 to 3 MJ/kg, a Reynolds number varying by a factor of four, and a leading edge geometry varied from sharp to one with a bevel of 0.2 mm. For the first time, the time dependence of flow macroparameters on the leading edge nose radius and the Reynolds number are studied using global linear theory. High-fidelity DSMC simulations showed that the temporal behavior of the separation region, which has significant effects on the surface parameters, depends closely on the leading edge bluntness and wall temperature. The formation of a secondary vortex was seen in about 2 ms for the sharp leading edge, whereas in the rounded leading edge geometry, it formed at earlier 0.7 ms. At a steady state, the size and structure of the separation zone, vortex structures, and surface parameters predicted by DSMC were found to be in good agreement with computational fluid dynamics for the higher density case. Finally, linear stability theory showed that for some leading edge shapes and flow densities, the time to reach the steady state was longer than the facility measurement time.

LIGS measurements in the nozzle reservoir of a free-piston shock tunnel

Free-piston shock tunnels are ground-based test facilities allowing the simulation of reentry flow conditions in a simple and cost-efficient way. For a better understanding of the processes occurring in a shock tunnel as well as for an optimal comparability of experimental data gained in shock tunnels to numerical simulations, it is highly desirable to have the best possible characterization of the generated test gas flows. This paper describes the final step of the development of a laser-induced grating spectroscopy (LIGS) system capable of measuring the temperature in the nozzle reservoir of a free-piston shock tunnel during tests: the successful adaptation of the measurement system to the shock tunnel. Preliminary measurements were taken with a high-speed camera and a LED lamp in order to investigate the optical transmissibility of the measurement volume during tests. The results helped to successfully measure LIGS signals in shock tube mode and shock tunnel mode in dry air seeded with NO. For the shock tube mode, six successful measurements for a shock Mach number of about 2.35 were taken in total, two of them behind the incoming shock (p $$\approx $$ 1 MPa, T $$\approx $$ 600 K) and four after the passing of the reflected shock (p $$\approx $$ 4 MPa, T $$\approx $$ 1000 K). For five of the six measurements, the derived temperatures were within a deviation range of $$6\%$$ to a reference value calculated from measured shock speed. The uncertainty estimated was less than or equal to $$3.5\%$$ for all six measurements. Two LIGS signals from measurements behind the reflected shock in shock tunnel mode were analyzed in detail. One of the signals allowed an unambiguous derivation of the temperature under the conditions of a shock with Mach 2.7 (p $$\approx $$ 5 MPa, T $$\approx $$ 1200 K, deviation $$0.5\%$$ , uncertainty $$4.9\%$$ ).

Shock-tunnel investigations on the evolution and morphology of shock-induced large separation bubbles

ABSTRACTShock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied bylargeseparation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

逆噴射ジェットによる高エンタルピー流中での空力加熱防御

An experimental investigation on an opposing jet as a thermal protection system in high enthalpy flow has been conducted. The heat flux distributions on a hemispherical-cylindrical test model with and without the opposing jet have been measured in a free-piston shock tunnel. The test gas is air with Mach number of 7.5 or 7.9 with total enthalpy of 20 MJ/kg or 4.8 MJ/kg, respectively. Nitrogen as a coolant gas is injected directly against the test flow through a sonic nozzle at the nose tip of the test model. The effect of alleviating heat loads is investigated by changing total pressure of the opposing jet. A significant reduction of surface heat loads is observed in the case of relatively strong opposing jet under both test flow conditions. The result indicates that thermal protection with the opposing jet is effective in high enthalpy flow. In addition, parameters to characterize the cooling effect of the opposing jet are discussed.

Numerical study on wall pressure over cone region of blunt-nosed body in high enthalpy shock tunnel HIEST

A cause of an overestimation of the computed surface pressure on a blunted cone in high-temperature hypersonic flow is explored. The overestimation was observed in a free-piston shock tunnel at the stagnation of H0=15.6 and 10.1 MJ/kg. The sensitivity analysis reveals that a reduction of the upstream translational temperature in the range of 100 to 300 K substantially improves the agreement of the surface pressure with the measured data. As the cause of the lower upstream translational temperature, radiative cooling effect is included in the thermochemical nonequilibrium calculation in the nozzle, the translational temperature at the nozzle exit is reduced to about 250 K. Using the obtained flow variables as the upstream boundary condition, the computed pressure agrees quite well with the experimental data. In order to clarify whether other variables such as translational–vibrational relaxation time, chemical reaction rates, and upstream chemical composition could be the cause of the discrepancy, uncertainty quantification is employed. It is shown that these parameters of the thermochemical model and upstream chemical composition have minor effect on the agreement of surface pressure. It is concluded that the observed discrepancy in the surface pressure is due to radiative cooling effect of high temperature gas in the nozzle region.

Rate-dependent energetic processes in hypersonic flows

In celebration of the first 60 years of the Air Force Office of Scientific Research, several studies of hypersonic flows dominated by rate-dependent energetic processes are revisited. The work presented shows the evolution and advancement of computational capabilities in this area, and illustrates some key lessons learned over the previous decade or so. Early work with Leyva and Hornung in the California Institute of Technology T5 Free-Piston Shock Tunnel had the goal of validating thermochemical models for high-enthalpy flows. Several of these flows are re-analyzed with more advanced numerical methods, resulting in improved comparisons with the experimental measurements. This work was followed by a series of experiments in the Calspan-University at Buffalo Research Center (now CUBRC Inc.) facilities at lower enthalpy conditions. Initial comparisons were poor, but with a better understanding of the facility behavior and the inclusion of key finite-rate processes, excellent agreement was obtained for nitrogen flows. An interesting study related to plasmadynamics and finite-rate processes in a different type of flow is discussed. Finally, it is shown that recent advances in numerical methods that are beginning to enable the direct numerical simulation of key rate-dependent energetic processes in hypersonic flows.

S1910103 高次精度スキームを用いた大気圏突入カプセル後方部の強制遷移流れ場計算([S191-01]大気突入・減速技術(1),宇宙工学部門)

An aeroheating measurement test campaign on an Apollo capsule model with turbulent boundary layers was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. When flow became turbulent flow, heat flux on fore and rear body was 1.5-2 times larger than one in laminar flow. In order to reproduce the heat flux by numerical simulation, we constructed high order CFD code towards LES in hypersonic flow. Since robustness near shock and high resolution in boundary layer is needed for LES in hypersonic flow, we employed recently improved WENO method. We calculated the hypersonic flow (M_∞ = 17) around cylinder and investigated the robustness for strong shock by this method. We could obtain smooth pressure distribution and good agreement with the calculated heat flux by NASA LAURA code. The hypersonic flow field around HTV-R, which is a manned space capsule under development by JAXA was also calculated towards predicting heat flux on after body.

極超音速乱流場解析に向けたWENO法に基づくLESコード開発

An aeroheating measurement test campaign on an HTV-R capsule model with turbulent boundary layers was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. When flow became turbulent flow, heat flux on rear body was 2-3 times larger than one in laminar flow. In order to reproduce the turbulent heat flux by numerical simulation, we constructed high order CFD code towards LES in hypersonic flow. Since robustness near shock and high resolution in boundary layer is needed for LES in hypersonic flow, we employed recently improved WENO method. We calculated the hypersonic flow (M_∞=17) around cylinder and investigated the robustness for strong shock by this method. We could obtain smooth pressure distribution and good agreement with the calculated heat flux by NASA LAURA code. The hypersonic flowfield around HTV-R was also calculated towards predicting turbulent heat flux on after body.

Modelling the complete operation of a free-piston shock tunnel for a low enthalpy condition

Only a limited number of free-stream flow properties can be measured in hypersonic impulse facilities at the nozzle exit. This poses challenges for experimenters when subsequently analysing experimental data obtained from these facilities. Typically in a reflected shock tunnel, a simple analysis that requires small amounts of computational resources is used to calculate quasi-steady gas properties. This simple analysis requires initial fill conditions and experimental measurements in analytical calculations of each major flow process, using forward coupling with minor corrections to include processes that are not directly modeled. However, this simplistic approach leads to an unknown level of discrepancy to the true flow properties. To explore the simple modelling techniques accuracy, this paper details the use of transient one and two-dimensional numerical simulations of a complete facility to obtain more refined free-stream flow properties from a free-piston reflected shock tunnel operating at low-enthalpy conditions. These calculations were verified by comparison to experimental data obtained from the facility. For the condition and facility investigated, the test conditions at nozzle exit produced with the simple modelling technique agree with the time and space averaged results from the complete facility calculations to within the accuracy of the experimental measurements.

0902 高エンタルピ衝撃風洞での境界層遷移トリップの開発(OS9-1 超音速流れの基礎と応用,OS9 超音速流れの基礎と応用)

Development of boundary layer trip for capsule shaped reentry vehicles under high-enthalpy hypersonic condition was developed in the free-piston shock tunnel HIEST. Wind tunnel test campaign with an 6% scaled Apollo CM model was conducted with stagnation pressure of 30MPa to 60Mpa and with stagnation enthalpy from 3.5MJ/kg to 22MJ/kg. Reynolds number based on model diameter was varied from 0.2 to 1.3 million. To measure heat flux around the model, eighty-four miniature co-axial thermocouples were instrumented on the heat shield surface of the model. Angle-of-attack of the model was fixed to 28degree. A pizza-box configuration boundary layer trip was mounted on the heat shield surface of the model, which trip height k were 0.3mm, 0.6mm and 0.8mm. With k=0.6mm height trip, boundary layer became fully turbulent less than H0=10MJ/kg. However, boundary layer was still laminar over H0=15MJ/kg even with k=0.8mm trip.

Development, Uniformity and Peak Heating of a Hypersonic Laminar Near-Wake Flow

This paper describes visualisation experiments performed using the planar laser-induced fluorescence of nitric oxide. The focus was to investigate the establishment and cross-stream uniformity of the hypersonic separated flow generated in a pulsed free-piston shock tunnel facility. In particular, the laminar Mach 7.5 flow over a step downstream of a wedge was investigated. The ratio of wedge width to total horizontal distance downstream of the leading edge was 0.71 at the furthest downstream measurement point, and model side plates were not used. Flow images were obtained on successive tunnel runs with different delays after flow onset, to identify the nominal flow establishment time. Furthermore, images were obtained at the nominal facility test time using a laser sheet oriented perpendicular to the step, propagating in the spanwise direction at a height equal to half the step height. These images are compared with thermocouple heat flux measurements obtained both with and without the step and show that ...

Improvement of a Model Scramjet Engine Design Using Various Types of Ground Tests

The scramjet, one of the most promising candidates for future transport systems, has many advantages such as a simple configuration and a high specific impulse. However, various characteristics of scramjet engines in a hypersonic environment are difficult to predict from theory. In the present study, we investigated the characteristics and the performance of the scramjet engine with various types of ground tests. An engine model (S1) was tested in the free-piston shock tunnel T4. (S1) showed active supersonic combustion and good starting characteristics, but its operating range is rather limited. To improve the performance of the model scramjet engine, we also investigated major components, such as the combustor and the intake by experiments with different model configurations. In the combustor test results, the zigzag cavity flame holder showed the best combustion performance among the three cases. By testing various intake configurations, the most efficient intake configuration was also determined. A revised model (S2), which was redesigned based on the component test results, showed enhanced thrust performance and a wider operating range in the free jet test using the free piston shock tunnel HIEST.

Cowl and Cavity Effects on Mixing and Combustion in Scramjet Engines

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel–air mixing.

Parametric Study of Stationary Body Facing its Own Wake in a High Speed Compressible Rotating Flow

Strongly rotating flow which is characterized by steep exponential density rise in the radial direction provides unique opportunity to analyze high speed flow for longer durations as compared to conventional shock tunnels, free-piston shock tunnels and expansion tubes. Cylinder wall speed and the position of the test specimen decide the value of upstream Mach number for such studies. Experimental investigations were performed to understand the effect of these two parameters on the stagnation pressure and temperatures of a stationary body placed in high speed rotating flow. A stationary body when placed in a rotating flow field faces its own wake and as a result produces slowdown. In this paper a radial entropy factor, Ref has been introduced to account for the entropy generation which not only depends on Mach number but also on the density of the flow which varies exponentially with respect to radial position. A simple mathematical model has also been developed to understand the complex effects of wall sp...

Force Measurement Techniques for Hypersonic Flows in Shock Tunnels

Recent advances in hypersonic research activities have led to the development in ground based test facilities for simulating near-realistic flight conditions in the laboratory. Basic performance data such as drag, lift, thrust etc., are usually obtained from ground-based experimentation for physical understanding of the flow fields and subsequent prototype developments. Hypersonic flows are produced for a very short time in impulse facilities such as shock tunnels, free-piston shock tunnels and expansion tubes. The test times are in the order of few milliseconds in the shock tunnel and for the expansion tubes, it is still less (~50 μs). The force measurements in these facilities are very challenging due the need of fast response sensors and instrumentations. In this review paper, two force measurement techniques are discussed that are best suited for short duration facilities namely; accelerometer balance and stress-wave force balance system.

Laminar Near Wake of a Circular Cylinder at Hypersonic Speeds

The laminar near-wake flow behind a circular cylinder at hypersonic speeds has been examined experimentally and analytically. Surface pressure and heat flux measurements were obtained in a free-piston shock tunnel at a nominal Mach number of 10. The freestream unit Reynolds numbers were 3.02 x 10 5 /m and 11.7 x 10 5 /m at total specific enthalpies of 13.35 and 3.94 MJ/kg, respectively. The experimental data of surface pressure and heat flux showed good agreement with theory based on perfect gas. Unlike surface pressure, the surface heat flux depended strongly on the wall-to-total-temperature ratio, and it increased with increase in the ratio. The surface pressure, however, depended on the Reynolds number, although this dependency was found to be weak. The flow separation angles behind the cylinder showed an inverse of the square root of Reynolds number dependence similar to the low-speed laminar flow separation behind a circular cylinder.

S1904-2-1 高圧・高エンタルピ流れにおけるApollo模型の空力過熱(大気突入・減速技術(2),社会変革を技術で廻す機械工学)

Aeroheating of an Apollo CM model was measured in the free-piston shock tunnel HIEST. The primary purpose of the measurement was to obtain benchmark data for numerical codes validation under high-enthalpy high-Reynolds number conditions. The model was made of SUS-304 stainless steel and had a diameter of 250mm. The windward surface of the model was equipped with 84 miniature co-axial Chromel-Constantan thermocouples that were manufactured in-house at JAXA. Twelve thermocouples and two Piezo-resistive pressure transducers were also mounted aft on the model and were used to determine the establishment of flow around the model. In this test campaign, heat flux data were measured for angles of attack 0° and 30°. Stagnation enthalpy and stagnation pressure were varied from H_0=4MJ/kg to 21MJ/kg, and from P_0=14MPa to 58MPa, respectively. Aeroheating characteristics were observed with a fully laminar boundary layer and with a transition boundary layer. Measured heat flux was normalized by the product of the Stanton number and the square-root of the Reynolds number for correlation with other wind tunnel results and flight results. The normalized heat flux value did not change remarkably when stagnation enthalpy was changed from H_0=4MJ/kg to 12MJ/kg, implying that the boundary layer flow was usually laminar. However, shots under high enthalpy and high Reynolds number conditions (H_0 over 18MJ/kg and P_0 approximately 50MPa) showed that the measured heat flux was approximately 40% higher than that obtained in other conditions.

Experimental investigation of hypersonic flow induced separation over double wedges

Flow separation occurs over the compression corners generated by deflected control surfaces on hypersonic re-entry vehicles and in the inlet of scram jet engines. Configurations like a double wedge and double cone model are useful for studying the separated flow features. Flow fields around concave corners are relatively complicated and produce several classical viscous flow features depending on the combination of the first and second wedge or cone half apex angles. Particularly characteristic phenomena are mainly shock/boundary layer, shock/shock interaction, unsteady shear layers and non-linear shock oscillations. Although most of these basic gas dynamics characteristics are well known, it is not clear what happens at high enthalpy conditions. This paper reports a result of flow fields over a double wedge at a stagnation enthalpy of 4.8 MJ/kg. The experiment was carried out in a free piston shock tunnel at a nominal Mach number of 6.99. Schlieren and double exposure holographic interferometry were applied to visualize the flow field over the double wedge.