Shock Tunnel Research Articles

S1910204 高温衝撃風洞HIESTによるHTV-Rカプセル形状模型の空力加熱試験([S191-02]大気突入・減速技術(2),宇宙工学部門)

Aerodynamic heating test of a HTV-R capsule model has performed in the free-piston high-enthalpy shock tunnel JAXA-HIEST. To measure heat flux distribution, the eighty-eight and thirty-four fast-response coaxial thermocouples were instrumented on the windward surface and on the back-shell surface, respectively. Besides, Eight Piezo-resistive pressure transducers were mounted on the aft of the body to determine flow establishment around the model through millisecond order short test period. Stagnation enthalpy and stagnation pressure were varied from H_0=3.5MJ/kg to 20MJ/kg, and from P_0=14MPa to 60MPa, respectively. The unit Reynolds number was changed from 0.9 million/m to 4.5 million/m. Aeroheating characteristics was observed at model angle of attack 20 degree with fully laminar, transition and turbulent boundary layer. Turbulent heat flux was observed with 'Pizza-box' shaped isolated roughness elements on the heat-shield surface and back-shell windward surface. Measured heat flux was normalized as Stanton number for comparison with different flow condition.

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.

Boundary-layer stabilization by an ultrasonically absorptive material on a cone in hypersonic flow: numerical investigations

The damping of acoustic second mode instabilities by passive porous surfaces is investigated numerically and compared with experiments for different hypersonic boundary-layer flows. The used geometry is a blunt 7° half-angle cone model with exchangeable nose. The cone surface consists partly of an ultrasonically absorptive material, which has a natural, random porosity. For the analyses of the second modes, the DLR stability code NOLOT, NOnLocal Transition analysis code, is used. The influence of the nose radius and the unit Reynolds number on the second modes is investigated using a smooth surface. These results of the smooth surface are compared with those of the porous surface to study the damping effect on the second modes and the transition shift. The numerical results are compared with wind tunnel measurements, which were performed in the DLR High Enthalpy Shock Tunnel Gottingen.

Experiments on passive hypersonic boundary layer control using ultrasonically absorptive carbon–carbon material with random microstructure

For the first time, the influence of ultrasonically absorptive carbon–carbon material on hypersonic laminar to turbulent boundary layer transition was investigated experimentally. A 7° half-angle blunted cone with a nose radius of 2.5 mm and a total length of 1,077 mm was tested at zero angle of attack in the High Enthalpy Shock Tunnel Gottingen of the German Aerospace Center (DLR) at Mach 7.5. One-third of the metallic model surface in circumferential direction was replaced by DLR in-house manufactured ultrasonically absorptive carbon–carbon material with random microstructure for passive transition control. The remaining model surface consisted of polished steel and served as reference surface. The model was equipped with coaxial thermocouples to determine the transition location by means of surface heat flux distribution. Flush-mounted piezoelectric fast-response pressure transducers were used to measure the pressure fluctuations in the boundary layer associated with second-mode instabilities. The free-stream unit Reynolds number was varied over a range of Rem = 1.5 × 106 m−1 to Rem = 6.4 × 106 m−1 at a stagnation enthalpy of h0 ≈ 3.2 MJ/kg and a wall temperature ratio of Tw/T0 ≈ 0.1. The present study revealed a clear damping of the second-mode instabilities and a delay of boundary layer transition along the ultrasonically absorptive carbon–carbon insert.

Experimental Investigation of Heat Fluxes in the Vicinity of Protuberances on a Flat Plate at Hypersonic Speeds

Heat transfer rates measured in front and to the side of a protrusion on an aluminum flat plate subjected to hypersonic flow at zero angle of attack are presented for two flow enthalpies of approximately 2 MJ/kg and 4.5 MJ/kg. Experiments were conducted in the hypersonic shock tunnel (HST2) and free piston driven HST3 at a freestream Mach number of 8. Heat transfer data was obtained for different geometries of the protrusion of a height of 4 mm, which is approximately the local boundary layer thickness. Comparatively high rates of heat transfer were obtained at regions of flow circulation in the separated region, with the hottest spot generally appearing in front of the protuberance. Experimental values showed moderate agreement with existing empirical correlations at higher enthalpy but not at all for the lower enthalpy condition, although the correlations were coined at enthalpy values nearer to the lower value. Schlieren visualization was also done to investigate the flow structures qualitatively.

高温衝撃風洞HIESTにおける極超音速熱空力研究の課題( 先進的空力技術 第2部:流れの制御と実験・計測 第21回)

高温衝撃風洞HIESTにおける極超音速熱空力研究の課題( 先進的空力技術 第2部:流れの制御と実験・計測 第21回)

Experimental Investigation of Hypersonic Flow and Plasma Aerodynamic Actuation Interaction

For hypersonic flow, it was found that the most effective plasma actuator is derived from an electromagnetic perturbation. An experimental study was performed between hypersonic flow and plasma aerodynamic actuation interaction in a hypersonic shock tunnel, in which a Mach number of 7 was reached. The plasma discharging characteristic was acquired in static flows. In a hypersonic flow, the flow field can affect the plasma discharging characteristics. DC discharging without magnetic force is unstable, and the discharge channel cannot be maintained. When there is a magnetic field, the energy consumption of the plasma source is approximately three to four times larger than that without a magnetic field, and at the same time plasma discharge can also affect the hypersonic flow field. Through schlieren pictures and pressure measurement, it was found that plasma discharging could induce shockwaves and change the total pressure and wall pressure of the flow field.

Unsteady Behaviors of a Hypersonic Inlet Caused by Throttling in Shock Tunnel

A two-dimensional hypersonic inlet/isolator model that exhibits self-starting characteristics is tested with different exit throttling ratios at a freestream Mach number of 5.9 in a shock tunnel. Various flow characteristics are observed and measured by applying simultaneous high-speed Schlieren imaging and surface pressure measurements. The results indicate that the backpressure generated by the throttling device can be tolerated, and the inlet can maintain the starting mode at low throttling ratios, whereas unstart flows are initiated from the near-choke throttling ratios, and a shock wave oscillation appears. The Schlieren movie demonstrates that the upstream-propagating shocks in the duct play important roles during the oscillation cycles and that the formations of the upstream-propagating shocks are related to the downstream-propagating compression waves/shock waves that encounter the throttling section. The frequency of the shock wave oscillation increases with increasing exit throttling ratios, primarily because of the acceleration of the upstream propagation.

超長試験時間デトネーション駆動衝撃波風洞@中国科学院力学研究所(北京・中国)( グローバルAERO 海外空力機関訪問・滞在レポート 第5回)

超長試験時間デトネーション駆動衝撃波風洞@中国科学院力学研究所(北京・中国)( グローバルAERO 海外空力機関訪問・滞在レポート 第5回)

Laser ignition of hypersonic air–hydrogen flow

An experimental investigation of the behaviour of laser-induced ignition in a hypersonic air–hydrogen flow is presented. A compression-ramp model with port-hole injection, fuelled with hydrogen gas, is used in the study. The experiments were conducted in the T-ADFA shock tunnel using a flow condition with a specific total enthalpy of 2.5 MJ/kg and a freestream velocity of 2 km/s. This study is the first comprehensive laser spark study in a hypersonic flow and demonstrates that laser-induced ignition at the fuel-injection site can be effective in terms of hydroxyl production. A semi-empirical method to estimate the conditions in the laser-heated gas kernel is presented in the paper. This method uses blast-wave theory together with an expansion-wave model to estimate the laser-heated gas conditions. The spatially averaged conditions found with this approach are matched to enthalpy curves generated using a standard chemical equilibrium code (NASA CEA). This allows us to account for differences that are introduced due to the idealised description of the blast wave, the isentropic expansion wave as well as thermochemical effects.

Prediction of short-duration transient surface heat flux using various analytical techniques

Transient surface heat flux prediction, from temperature signals using one-dimensional heat conduction modeling, is the major objective of present investigation. The techniques reported in the literature based on spline fitting (linear and cubic) and least square polynomial fitting of temperature data are evaluated for prediction of surface heat flux through various analytical modeling. In addition, a Laplace-based technique is also incorporated here to predict surface heat flux where the least square polynomial fitting technique is used to discretize the temperature data. The temperature time histories obtained from an in-house built, one-dimensional finite volume computation solver, and experiments (i.e., shock tunnel testing and flight data) are considered for the performance assessment of these methods. Heat flux recovery from all the methods for smooth temperature signals is seen to be in good agreement with a reasonable accuracy of ±5%. However, it has been noticed that the spline based fitting techniques supersede the polynomial-based fitting techniques for prediction of heat flux from discontinuous or noisy temperature signals. © 2013 Wiley Periodicals, Inc. Heat Trans Asian Res; Published online in Wiley Online Library (wileyonlinelibrary.com/journal/htj). DOI 10.1002/htj.21050

Design and testing of an external drag balance for a hypersonic shock tunnel

An external force balance for a hypersonic shock tunnel was developed. The design utilized finite element analysis that identified the dynamics of the balance. A simulated impulse and a step load were applied to the design, the former for determining the simulated transfer function and the latter to validate the design. The numerical modeling showed the feasibility of this approach for designing stress wave force balances. A force balance based on the design was fabricated, calibrated statically and dynamically, and implemented in a shock tunnel for measuring drag of a blunt cone at Mach 9.4. The measured drag compared well with modified Newtonian theory.

Hypersonic Experimental Analysis of Impulse Generation in Airbreathing Laser Thermal Propulsion

This experimental campaign focused on the analysis of the feasibility for impulse generation of a repetitively pulsed airbreathing laser lightcraft at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of nano- and microsatellites (i.e., 1–100 kg payloads) into low Earth orbit, at low cost and on demand. These laser propulsion experiments employed a hypersonic shock tunnel, integrated with twin gigawatt pulsed Lumonics 620-TEA lasers to produce the required test conditions. Time-dependent surface pressure distributions were recorded together with schlieren movies of the flowfield structure resulting from the laser energy deposition, at nominal Mach numbers ranging from 6 to 10. The laser induced breakdown occurring off-surface and across the inlet’s midchannel was analyzed and discussed against previous theoretical and experimental results. The data indicated laser induced pressure increases of 0.7–0.9 bar with laser pulse energies of , on an off-shroud induced breakdown condition and a freestream Mach number of six. The results of this research corroborate the feasibility of laser powered airbreathing flight with infinite specific impulse ().

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.

Computational and experimental analysis of Mach 2 air flow over a blunt body with plasma aerodynamic actuation

According to the mechanism of the arc plasma heating effect, and from a phenomenological perspective of view, the plasma actuation was simplified as heating energy injected into the supersonic flow field for the numerical research on controlling detached shock of the blunt body in non-center symmetrical positions. Besides, experimental research on the form and strength of detached shock wave control by plasma aerodynamic actuation in non-center symmetrical positions was conducted in a high-speed shock tunnel (M=2). The results showed that the detached distance of shock wave increased and the strength of normal shock wave ahead of the detached shock wave reduced when plasma actuation was applied. The control effect was greatly improved after the magnetic field was applied and the effect of upwind-direction flow was the best one. When the upwind-direction flow was applied with 1000 V voltage actuation, the distance of detached shock wave would increase from 3.4 to 7.6 mm and the time average strength of normal shock wave was weaken by 5.5%. At last, the mechanism of plasma actuation on controlling the detached shock wave was briefly analyzed.

Interaction of Laser Energy with Bow Shock in Mach 5 Flow

In order to develop a potential way to reduce the stagnation pressure and heat flux of the blunt body in hypersonic flow, single pulsed and high rated laser energy are deposited respectively to control the bow shock in Mach 5 flow. Process of the interaction of 100mJ single pulsed laser energy with bow shock is studied by schlieren experiment in Mach 5 shock tunnel. The results indicate that the schlieren photographs and stagnation pressure at different times fit well with the simulation. Bow shock is distorted by the laser induced blast wave. During a single laser pulse, the average stagnation pressure and temperature are reduced by 6.5% and 3.4%. High rated laser energy is used to sustain the low pressure and heat flux region to increase the control efficiency. With the power of 6.6% of the enthalpy flux, when the distance of deposition point and blunt body is 1.1 times of the diameter of the blunt body, the pressure and heat flux of the stagnation point is reduced to 49% and 75%, respectively. The mechanism of interaction of laser and bow shock is disclosed.

Stabilization of a Hypersonic Boundary Layer Using a Wavy Surface

Stability of a supersonic near-wall flow over a shallow grooved plate in the freestream of Mach 6 is investigated by means of numerical simulations and wind-tunnel experiments. Numerical solutions of two-dimensional Navier–Stokes equations are used to model propagation of artificial disturbances of several fixed frequencies generated by an actuator placed on the wall. It is shown that the high-frequency forcing excites unstable waves in the flat-plate boundary layer. These waves are relevant to the second-mode instability. The wavy wall damps the disturbances in a high-frequency band while it enhances them at lower frequencies. Stability experiments are conducted in the Institute of Theoretical and Applied Mechanics Tranzit-M shock tunnel under natural freestream conditions. The measured disturbance spectra are similar to those predicted numerically. They contain a peak associated with the second-mode instability. This peak is damped by the wavy wall, while a marginal increase of the disturbance amplitude is observed at lower frequencies. Although the experiments qualitatively confirm the wavy-wall stabilization concept, further stability and transition measurements are needed to clarify its robustness.

Design of a High Resolved Schlieren System

A schlieren system with high time and space resolution was developed to diagnose the hypersonic flow field with high speed, low density and intensive shock interactions. The key technique of color filter used for color schlieren was developed. The time resolution reached to 100ns and the space resolution reached to 0.1mm. In shock tunnel with Mach 5, bow shock controlled by single laser pulse was investigated with the schlieren system, and the interaction of oblique shocks generated by double ramps was detected with color schlieren. The complex flow fields are clearly exploded, which proves that the schlieren system designed in this paper supplies an efficient method in the diagnostics of hypersonic flow field.

Pressure, temperature and free-flight projectile-displacement measurements during a plasma discharge generated on a supersonic projectile

The generation of a plasma discharge on a projectile surface is a way of producing a pressure imbalance in order to divert a projectile from its initial trajectory. Thus, some experiments dealing with the steering of a supersonic projectile by using a plasma actuator are presented. The plasma discharge is produced by an embedded low-voltage generator capable of delivering a power controlled electric discharge between the electrodes flush with the projectile surface. A series of experiments carried out under wind-tunnel conditions at a Mach number of 3 allows surface-pressure and temperature measurements, which are synchronised with the visualisation of the perturbation evolution along the projectile surface, the disturbance being generated by the plasma discharge. Other series of experiments realised in a shock tunnel under 2.5 km of altitude real conditions at a Mach number of 4.5 demonstrates that a plasma discharge generates the angular deviation of an explosively-formed-projectile model: 120 J of stored energy delivered in 4.5 ms in front of the ogive-cylinder junction of the projectile induces an angle of attack of 11° after the plasma-discharge generation.