Shock Tunnel Research Articles

Improved quantitative visualization of hypervelocity flow through wavefront estimation based on shadow casting of sinusoidal gratings.

A simple noninterferometric optical probe is developed to estimate wavefront distortion suffered by a plane wave in its passage through density variations in a hypersonic flow obstructed by a test model in a typical shock tunnel. The probe has a plane light wave trans-illuminating the flow and casting a shadow of a continuous-tone sinusoidal grating. Through a geometrical optics, eikonal approximation to the distorted wavefront, a bilinear approximation to it is related to the location-dependent shift (distortion) suffered by the grating, which can be read out space-continuously from the projected grating image. The processing of the grating shadow is done through an efficient Fourier fringe analysis scheme, either with a windowed or global Fourier transform (WFT and FT). For comparison, wavefront slopes are also estimated from shadows of random-dot patterns, processed through cross correlation. The measured slopes are suitably unwrapped by using a discrete cosine transform (DCT)-based phase unwrapping procedure, and also through iterative procedures. The unwrapped phase information is used in an iterative scheme, for a full quantitative recovery of density distribution in the shock around the model, through refraction tomographic inversion. Hypersonic flow field parameters around a missile-shaped body at a free-stream Mach number of ∼8 measured using this technique are compared with the numerically estimated values. It is shown that, while processing a wavefront with small space-bandwidth product (SBP) the FT inversion gave accurate results with computational efficiency; computation-intensive WFT was needed for similar results when dealing with larger SBP wavefronts.

Ignition Experiments of Hydrocarbons in a Mach 8 Shape-Transitioning Scramjet Engine

The design and testing of a cavity flameholder in a Mach 8 rectangular-to-elliptical shape-transitioning scramjet is presented. The experimental results presented are from the shock tunnel testing of the engine, fitted with the cavity combustor and fueled by gaseous ethylene, methane, and hydrogen. Axisymmetric simulations of a cavity in a diverging combustor were used to confirm that the cavity flow would fully establish in the limited test time available in our impulse facility. The cavity flameholder was designed such that autoignition and flameholding of gaseous ethylene–air mixtures should occur. Static pressure measurements were taken along the entire flowpath of engine. The engine was fueled at equivalence ratio ranging from . Successful combustion and flameholding of ethylene and hydrogen was observed at a flight Mach number of 7.3 and a dynamic pressure of 53.5 kPa at an altitude of 28.7 km. Methane did not ignite at any of the conditions tested.

Experiments on High-Temperature Xenon Plasma Magnetohydrodynamic Power Generation

Magnetohydrodynamic power generation experiments with seed-free pure xenon plasma have been conducted using a shock tunnel. The effects of inlet total temperatures between 5000 and 11 500 K and magnetic flux density on the generator performance and plasma characteristics are revealed. An enthalpy extraction ratio of 4.1% is achieved at an inlet total temperature of 6000 K, which is higher than that for argon at the same inlet total temperature. The enthalpy extraction ratio increases with increasing inlet total temperature and magnetic flux density, but saturates at a high inlet total temperature (>8000 K). The inhomogeneous and unstable plasma due to ionization instability at a low inlet total temperature changes to a homogeneous and stable state with an increase in the inlet total temperature. The change in plasma structure occurs at a temperature between 6500 and 7000 K, which is lower than that in the case of argon. Moreover, plasma stabilization is achieved at a low magnetic flux density even when the inlet total temperature is low.

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.

충격파 풍동에서의 자유 낙하 장치를 활용한 힘 측정

In this paper, acceleration and pressure exerted on a human model were measured under a supersonic condition in a shock tunnel. In order to measure these in an interference-free environment, free-fall technique with an electromagnet and a three-dimensional iron-powdered human model was used. Free-fall experiment was conducted at Mach 4 and the force acting on the model was obtained by calculating the displacement from the flow visualization images.

Force measurement using strain-gauge balance in a shock tunnel with long test duration.

Force tests were conducted at the long-duration-test shock tunnel JF12, which has been designed and built in the Institute of Mechanics, Chinese Academy of Sciences. The performance tests demonstrated that this facility is capable of reproducing a flow of dry air at Mach numbers from 5 to 9 at more than 100 ms test duration. Therefore, the traditional internal strain-gauge balance was considered for the force tests use in this large impulse facility. However, when the force tests are conducted in a shock tunnel, the inertial forces lead to low-frequency vibrations of the test model and its motion cannot be addressed through digital filtering because a sufficient number of cycles cannot be found during a shock tunnel run. The post-processing of the balance signal thus becomes extremely difficult when an averaging method is employed. Therefore, the force measurement encounters many problems in an impulse facility, particularly for large and heavy models. The objective of the present study is to develop pulse-type sting balance by using a strain-gauge sensor that can be applied in the force measurement of 100 ms test time, especially for the force test of the large-scale model. Different structures of the S-series (i.e., sting shaped balances) strain-gauge balance are proposed and designed, and the measuring elements are further optimized to overcome the difficulties encountered during the measurement of aerodynamic force in a shock tunnel. In addition, the force tests were conducted using two large-scale test models in JF12 and the S-series strain-gauge balances show good performance in the force measurements during the 100 ms test time.

Fore-body and base heat flux measurements on a typical crew module in short duration impulse facilities

Experiments are carried out to measure forebody and base heat flux on a typical Crew Module at Mach numbers 5.8 and 8 over the enthalpy range of 1.5–3.73MJ/kg. Towards this, two scaled down models namely 1:55 and 1:40 are tested in 0.15m pressure and in 1m combustion driven shock tunnels. Forebody and stagnation point heat flux are measured through E-type co-axial thermocouples whereas, platinum thin film gauges are used to measure base heat flux. Forebody heat flux for the range of α (0–20deg) shows that the stagnation point heat flux move towards the bottom surface and the top surface experiences low heat flux as the angle of attack increases. At the base, the level of heat flux is found to be low and the maximum value is 3% of the stagnation point heat flux. The minimum heat flux is observed to be at the center of the base and increases radially outwards. The measured stagnation point heat flux is found to be ∼50% higher than the Fay–Riddell correlation for dissociated air, using Newtonian velocity gradient. On the other hand, if the Fay–Riddell correlation is modified by incorporating experimentally generated velocity gradient, the difference reduces to 14.25%. Finally, once the real gas effect is added to the Fay–Riddell correlation along with experimentally generated velocity gradient, the difference further reduces to 9.4%. Thus, the correct choice of transport properties and velocity gradient plays a crucial role to accurately predict the stagnation point heat flux.

Starting Process in a Large-Scale Shock Tunnel

The starting process of the flow in an expansion nozzle (nominal Mach number 6) with an outlet diameter of 1.5 m and 8.9 m length, which is used for a large-scale hypersonic shock tunnel in the Key Laboratory of High-Temperature Gas Dynamics, was simulated and analyzed at incident Mach number . The calculating domains include the driven section (the shock-tube end, about 4.8 m length), the nozzle with about 8.9 m length, and part of the test section (more than 3 m). The characteristics of unsteady nozzle flow, including the shock wave patterns in the nozzle inlet region and inside the nozzle, were analyzed numerically in the viscous and inviscid flow regimes, respectively. The pressure and Mach number results were presented and discussed by comparing with the experimental findings, where the simulated results of the reflected shock wave in the shock tube and the transmitted shock wave inside the nozzle were found to agree well with the test data. Additionally, the case without a contraction section for the throat configuration was also calculated and compared with the case with a contraction section. The effect of the starting flow in these two cases on the flowfield uniformity is discussed in detail in this paper.

Research on electron density measurements for high enthalpy flows

The high enthalpy shock tunnel, which can provide the high temperature gas conditions for hypersonic flight, is an effective ground facility for the research of reentry problems. It also has the ability of studying real gas effect. The hypersonic test flow condition, with the stagnation enthalpy of 16 MJ/kg and stagnation temperature of 7900 K, was achieved in the JF-10 high enthalpy shock tunnel. The electron density of the flow-field for the 10 degree wedge model was measured using the electrostatic probe, which can obtain sufficient spatial resolution without influencing the flow structure. Results showed that the electrostatic probe has high sensitivity and the spatial electron density can be preferable acquired. The constant bias method can get the dimensionless electron density coupling with the flow field parameters, but the new developed sweeping circuit can get the quantitative value on the basis of effectively reducing the induced noise in the circuit.

Comparative study on aerodynamic heating under perfect and nonequilibrium hypersonic flows

In this study, comparative heat flux measurements for a sharp cone model were conducted by utilizing a high enthalpy shock tunnel JF-10 and a large-scale shock tunnel JF-12, responsible for providing nonequilibrium and perfect gas flows, respectively. Experiments were performed at the Key Laboratory of High Temperature Gas Dynamics (LHD), Institute of Mechanics, Chinese Academy of Sciences. Corresponding numerical simulations were also conducted in effort to better understand the phenomena accompanying in these experiments. By assessing the consistency and accuracy of all the data gathered during this study, a detailed comparison of sharp cone heat transfer under a totally different kind of freestream conditions was build and analyzed. One specific parameter, defined as the product of the Stanton number and the square root of the Reynold number, was found to be more characteristic for the aerodynamic heating phenomena encountered in hypersonic flight. Adequate use of said parameter practically eliminates the variability caused by the deferent flow conditions, regardless of whether the flow is in dissociation or the boundary condition is catalytic. Essentially, the parameter identified in this study reduces the amount of ground experimental data necessary and eases data extrapolation to flight.

Design of a pulse-type strain gauge balance for a long-test-duration hypersonic shock tunnel

When the measurement of aerodynamic forces is conducted in a hypersonic shock tunnel, the inertial forces lead to low-frequency vibrations of the model, and its motion cannot be addressed through digital filtering because a sufficient number of cycles cannot be obtained during a tunnel run. This finding implies restrictions on the model size and mass as the natural frequencies are inversely proportional to the length scale of the model. Therefore, the force measurement still has many problems, particularly for large and heavy models. Different structures of a strain gauge balance (SGB) are proposed and designed, and the measurement element is further optimized to overcome the difficulties encountered during the measurement of aerodynamic forces in a shock tunnel. The motivation for this study is to assess the structural performance of the SGB used in a long-test-duration JF12 hypersonic shock tunnel, which has more than 100 ms of test time. Force tests were conducted for a large-scale cone with a $$10^{\circ }$$ semivertex angle and a length of 0.75 m in the JF12 long-test-duration shock tunnel. The finite element method was used for the analysis of the vibrational characteristics of the Model-Balance-Sting System (MBSS) to ensure a sufficient number of cycles, particularly for the axial force signal during a shock tunnel run. The higher-stiffness SGB used in the test shows good performance, wherein the frequency of the MBSS increases because of the stiff construction of the balance. The experimental results are compared with the data obtained in another wind tunnel and exhibit good agreement at $$M = 7$$ and $$\alpha =5^\circ $$ .

Base Flow of Circular Cylinder at Hypersonic Speeds

The paper presents a computational and an experimental investigation of base flow of a circular cylinder at hypersonic speeds. Effects of chemistry and wall temperature on the flow in the base region, at low to high enthalpies, are discussed. The experiments were conducted in a shock tunnel at a nominal Mach number of 10. Freestream Reynolds numbers based on cylinder diameter were and , respectively, and the total specific enthalpies were 13.35 and , respectively. The test gas was air. The surface pressure and heat flux were measured using a cold wall model. Equilibrium and thermal as well as chemical nonequilibrium numerical simulations were performed using a Navier–Stokes equations-based computational fluid dynamics code. Both a cold wall and adiabatic wall were considered. Particular emphasis was placed on the wake structure, vorticity distribution, wake centerline aerothermodynamic properties, and surface data. The existing low-enthalpy cold hypersonic wind-tunnel experimental data are included for comparison. The simulations predicted the effect of chemistry on the near wake to be negligible for the low-enthalpy, high Reynolds number flow but more significant for the high-enthalpy, low Reynolds number flow.

衝撃風洞によるリフティングボディ形状模型の空力特性評価

衝撃風洞によるリフティングボディ形状模型の空力特性評価

A gasdynamic gun driven by gaseous detonation.

A gasdynamic gun driven by gaseous detonation was developed to address the disadvantages of the insufficient driving capability of high-pressure gas and the constraints of gunpowder. The performance of this gasdynamic gun was investigated through experiments and numerical simulations. Much more powerful launching capability was achieved by this gun relative to a conventional high-pressure gas gun, owing to the use of the chemical energy of the driver gas. To achieve the same launching condition, the initial pressure required for this gun was an order of magnitude lower than that for a gun driven by high-pressure H2. Because of the presence of the detonation, however, a more complex internal ballistic process of this gun was observed. Acceleration of projectiles for this gun was accompanied by a series of impulse loads, in contrast with the smooth acceleration for a conventional one, which indicates that this gun should be used conditionally. The practical feasibility of this gun was verified by experiments. The experiments demonstrated the convenience of taking advantage of the techniques developed for detonation-driven shock tubes and tunnels.

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.

Experiments in hand-operated, hypersonic shock tunnel facility

Experiments were conducted using the newly developed table-top, hand-operated hypersonic shock tunnel, otherwise known as the Reddy hypersonic shock tunnel. This novel instrument uses only manual force to generate the shock wave in the shock tube, and is designed to generate a freestream flow of Mach 6.5 in the test section. The flow was characterized using stagnation point pressure measurements made using fast-acting piezoelectric transducers. Schlieren visualization was also carried out to capture the bow shock in front of a hemispherical body placed in the flow. Freestream Mach numbers estimated at various points in the test section showed that for a minimum diameter of 46 mm within the test section, the value did not vary by more than 3 % along any cross-sectional plane. The results of the experiments presented here indicate that the device may be successfully employed for basic hypersonic research activities at the university level.

Fast response co-axial thermocouple for short duration impulse facilities

A fast response Chromel–Alumel (K-type) co-axial thermocouple is designed, fabricated, calibrated and tested in a shock tunnel. The freestream Mach number of 5.75 and the total enthalpy of 0.92 MJ/kg are simulated to study the stagnation point heat flux of a hemi-spherical model through transient temperature trace. The realized K-type co-axial thermocouple of 3 mm in length and 1.6 mm in diameter is flush mounted at the stagnation point of a 7.5 mm radius hemi-spherical model. The achieved response time of the realized K-type co-axial thermocouple is ~3 µs, which is sufficient enough to capture the transient temperature signal. A steady tunnel flow time of 1.8 milliseconds is used to get the average stagnation point heat flux. The measured stagnation point heat flux is 22.96 W/cm2, which is well matched with the Fay–Riddell value within 5.5%. The realized K-type co-axial thermocouple is robust, responds fast, and can be contoured to any type of model surface.

Thermal and inertial parameters of the flow field of a scramjet engine

Heat flux and pressure were measured on the internal walls of a scramjet engine at a hypersonic Mach number of 8 in a shock tunnel. Temperature-time history was recorded on model walls using fast-response thermocouples and the wall heat flux was deduced thereon from the temperature records. Pressure was measured at the same locations as that of heat flux using high-frequency pressure transducers. The surface measurements were used to analyze the internal flow field of the engine. The heat flux measurements revealed the regions of peak heating that may require additional thermal protection during the operation; the regions being the beginning of the constant area duct and the nozzle. The pressure measurements and the pertinent analysis indicated an optimum compression at intake through shock waves with a viable flow separation in the combustion chamber. The experimental data indicated an optimum design of the engine for a possible fuelled operation in future.

Observations of hypervelocity boundary-layer instability

A novel optical method is used to measure the high-frequency (up to 3 MHz) density fluctuations that precede transition to turbulence within a laminar boundary layer in a hypervelocity flow. This optical method, focused laser differential interferometry, enables measurements of short-wavelength, high-frequency disturbances that are impossible with conventional instrumentation such as pressure transducers or hot wires. In this work, the T5 reflected-shock tunnel is used to generate flows in air, nitrogen and carbon dioxide with speeds between 3.5 and $5~\text{km}~\text{s}^{-1}$ (Mach numbers between 4 and 6) over a 5° half-angle cone at zero angle of attack. Simultaneous measurements are made at two locations approximately midway along a generator of the 1-m-long cone. With increasing Reynolds number (unit values were between 2 and $5\times 10^{6}~\text{m}^{-1}$), density fluctuations are observed to grow in amplitude and transition from a single narrow band of frequencies consistent with the Mack or second mode of boundary-layer instability to bursts of large-amplitude and spectrally broad disturbances that appear to be precursors of turbulent spots. Disturbances that are sufficiently small in initial amplitude have a wavepacket-like signature and are observed to grow in amplitude between the upstream and downstream measurement locations. A cross-correlation analysis indicates propagation of wavepackets at speeds close to the edge velocity. The free stream flow created by the shock tunnel and the resulting boundary layer on the cone are computed, accounting for chemical and vibrational non-equilibrium processes. Using this base flow, local linear and parabolized stability (PSE) analyses are carried out and compared with the experimental results. Reasonable agreement is found between measured and predicted most unstable frequencies, with the greatest differences being approximately 15 %. The scaling of the observed frequency with the inverse of boundary-layer thickness and directly with the flow velocity are consistent with the characteristics of Mack’s second mode, as well as results of previous researchers on hypersonic boundary layers.