Shock Tunnel Research Articles

Time-Resolved On-Axis Spectroscopic Stagnation Temperature Measurements in Shock Tunnel Flows

This paper demonstrates time-resolved stagnation temperature measurements in a shock tunnel at a frequency of 25 kHz using emission spectroscopy in air and nitrogen test conditions. The two most important parameters for determining the flow conditions generated in a shock tunnel experiment are the stagnation pressure and temperature of the flow just upstream of the supersonic nozzle. While the pressure can be measured using a wall-mounted transducer with relative ease, the measurement of the temperature requires a optical technique such as time resolved emission spectroscopy. Knowledge of the transient stagnation temperature behavior is critical to all subsequent expansion tube flow processes. The driver gas emission spectrum data at the post-shock condition shows continuum and atomic line radiation. The continuum radiation can be described by a black body radiator with the individual spectra showing sufficient continuum information for accurately fitting Planck functions. Atomic line radiation was excluded by skipping those data from the measured spectra. The fitting routine shows clear differences in determined temperatures including and neglecting atomic line radiation. These measurements allow for the exact determination of the shock tunnel flow conditions in combination with pressure transducer data. The flow condition used in the experiment corresponds to a nominal Mach-10 condition at an altitude of 65 km, however, the technique is not limited to this condition and can be used for a large range of flow conditions.

An accelerometer balance for aerodynamic force measurements over Hypervelocity Ballistic models in shock tunnel

A force balance to measure drag, lift and pitching moment over Hypervelocity Ballistic (HB) models 1 and 2 has been developed and assessed in an impulse test facility, hypersonic shock tunnel. The experiments were carried out at different angles of incidence of the models with the freestream at nominal Mach numbers of 5.78 and 8.31. The force balance is provided with rubber supports to ensure unrestrained motion of the model during the hypersonic freestream flow. Three dimensional finite element simulations are carried out to assess the performance of the force balance system in measuring the acceleration within the short test duration available in the shock tunnel. The force balance is mounted inside the model and is equipped with accelerometers to sense the model’s acceleration in the desired direction. The aerodynamic force and moment coefficients are then deduced from the measured acceleration signals based on theories in aerodynamics. The measured values of force and moment coefficients match well with the theoretical values estimated using modified Newtonian theory for HB 1 and HB 2 models. In addition, CFD simulations are carried out for the experimental test cases using viscous, laminar models and the numerically computed force coefficients are validated by comparing against the experimental values. The present study demonstrates the use of accelerometer based force balance for measuring aerodynamic forces and moments over realistic long missile configurations.

Experimental Investigation of Roughness Effects on Transition on Blunt Spherical Capsule Shapes

Boundary-layer transition plays an important role in aerodynamic heating of blunt capsule shapes. Transition may occur from disturbances created by surface roughness and fluctuations of the freestream. The present paper reports on a number of related experiments on Apollo-type capsule models in a range of hypersonic wind tunnels. The Adjustable Contour Expansion facility of the Texas A&M University provided Mach 6 experiments with distributed roughness at relatively low Reynolds numbers, , with denoting the capsule diameter. Observed boundary-layer transition data align well with current estimates of transient growth theory. Larger Reynolds numbers, , could be assessed in the hypersonic Ludwieg tube Braunschweig of the Braunschweig University of Technology in Germany. Distributed roughness height values were broadly varied during these experiments, and different fluctuation levels of the freestream were considered. The results indicate roughness-induced transition for larger roughness heights, in accordance with correlations based on transient growth theory, and transition due to freestream disturbances, at subcritical roughness. Effects of total flow enthalpy on transition with isolated roughness were investigated in the High Enthalpy Shock Tunnel facility of the Japan Aerospace Exploration Agency. Here, roughness height, Reynolds number, and flow total enthalpy were varied. The paper reviews the most important results of these experiments and provides analysis, taking into account transient growth theory and numerical simulation data.

High-Enthalpy Effects on Hypersonic Boundary-Layer Transition

In the present study, three boundary-layer stability codes are compared based on hypersonic high-enthalpy boundary-layer flows around a blunted 7 deg half-angle cone. The code-to-code comparison is conducted between the following codes: the Nonlocal Transition analysis code of the DLR, German Aerospace Center (DLR); the Stability and Transition Analysis for hypersonic Boundary Layers code of VirtusAero LLC; and the VKI Extensible Stability and Transition Analysis code of the von Kármán Institute for Fluid Dynamics. The comparison focuses on the role of real-gas effects on the second-mode instability, in particular the disturbance frequency, and deals with the question on how far not accounting for real-gas effects compromises the stability analysis. The experimental test cases for the comparison are provided by the DLR High Enthalpy Shock Tunnel Göttingen and the Japan Aerospace Exploration Agency High Enthalpy Shock Tunnel. The focus of the comparison between the stability results and the measurements is, besides real-gas effects, the influence of uncertainties in the mean flow on the stability analysis.

Weighting by Cross-Validation: A Calibration Method for Force Measurements via Transient Response Analysis

Transient response analysis is an indirect method for the measurement of the aerodynamic force experienced by hypersonic vehicles in impulse wind-tunnel facilities; a transfer function is identified by a calibration experiment in advance and is then used to recover the target aerodynamic force. Theoretically, the transfer function is unique for a given measurement system. However, the calibration experiment may involve unpredictable factors and noise that are inevitable in practical applications, which will result in systematic errors. In this paper, a new calibration method, weighting by cross-validation (WCV), is proposed to reduce the systematic errors. In WCV, a series of on-site calibration experiments are carried out to obtain a set of transfer functions prior to the wind tunnel test. The transfer functions are cross-validated against each other to create a cross-validation table of relative measurement errors. The aerodynamic force is then calibrated using the average and standard deviation of the cross-validation errors. The working mechanism of the WCV method is demonstrated by analogy of measurements with a set of non-standard rulers. The effectiveness of the WCV method has been verified by tests in the JF-12 shock tunnel. Studies show that the WCV method improves the measurement accuracy significantly. In addition, the concept of the WCV method is general and can also be applied to other indirect measurement problems to reduce systematic errors, especially when no exact/standard measurement tools are available.

Experimental investigation of heat transfer over double disk spike-blunt body at Mach 5.7

The forefront research in the territory to the regime of hypersonic flow is intended to focus on improving the overall performance of the existing aerospace vehicle. Two major hindrances for the development of an effective high speed vehicle are aerodynamic drag acts against the vehicle and heat transfer rate on the body. In this study, we explore the possibility of addressing this issue by placing an aerospike at the nose portion (i.e., stagnation region) of the high speed vehicle and as a result, wave drag of vehicle is considerably reduced. However, this will lead to the increase in heat transfer at the localized spot of the main body. In an attempt to reduce this localized high heat transfer, the current study evaluates the variation in the heat transfer on the body, by modifying the spike through the addition of a smaller hemisphere on the mid-section of the aerodisk spike. In this regard, shock tunnel experiments and computational studies were carried out on this modified spike configuration, termed as “double-disk spike” or “double spike”. The experimental results indicate that heat transfer near the localized spot of the blunt body with spike decreases for a double spike case in comparison to single disk spike. The decrease in heat transfer varies from 5% to 30% depending on the double spike configuration (i.e. varying cap radius and length of the spike). To supplement these results, 3D Finite Volume simulation, “HiFUN” (High Resolution Flow Solver on Unstructured), showed a good trend in heat transfer compared to that obtained from the experiment; it was found that simulated drag from the double spike is less than single spike case.

An innovative approach for prediction of aerodynamic coefficients in shock tunnel testing with soft computing techniques

A blunt bi-conic aluminum model, integrated with a three component accelerometer force balance system, is tested in the IITB-Shock Tunnel at various angle of inclinations with an intention to measure aerodynamic coefficients. Initially, Genetic Algorithm (GA) is employed to deduce the orthogonal inputs and their responses from distributed point loads applied on the test model during calibration experiments. Time histories of these loads and their responses are then used to train the architecture of Adaptive Neuro Fuzzy Inference System, ANFIS. This ANFIS architecture is further employed for force prediction from the acquired acceleration responses during shock tunnel experiments. Testing of the experimental acceleration responses of 0° and 10° angle of attack experiments showed encouraging agreement for recovered aerodynamic coefficients with the accelerometer balance theory based predictions. These results clearly showed the necessity to consider multiple point loads and their acceleration responses for training the soft computing algorithm or calibration of the force balance. Further, use of only one point force and its responses, for training ANFIS, is not only seen to have discrepancy in prediction but also is noted to be non-unique due to choice of the loading point. Moreover, it is recommended to choose the loading point near the center of pressure for the tested experimental conditions to incur lower discrepancy in prediction using single point loading data for ANFIS training.

Unsteady shock interactions on V-shaped blunt leading edges

The unsteady dynamics of shock interactions on the crotch of two typical V-shaped blunt leading edges have been investigated numerically and experimentally with a freestream Mach number 6 for different ratios (R), defined as the rounding radius at the crotch to the blunt radius at the leading edge. The primary flow features observed in the shock tunnel experiments are reproduced by the large-eddy simulations. The time-averaged flow structures in the crotch are clearly shown as counterrotating vortices originating from the collision of jets near the stagnation point. These jets and vortices undergo unsteady motions coupled with the dynamics of shock interactions. Of great interest, two typical global oscillations, i.e., swing oscillation and arch-recover oscillation corresponding to the two values of R, are identified. The coherent structures of the oscillations are analyzed using the proper orthogonal decomposition technique. It is demonstrated that the swing oscillation and arch-recover oscillation are characterized by an antisymmetric pattern and a symmetrical pattern, respectively. These two oscillations are also characterized by the energetic middle-frequency components of the broadband wall pressure spectra. Two feedback models are proposed for the prediction of such middle-frequency components. The results show that the swing oscillation causes much more severe pressure load, and the local impingement of the transmitted shock on the crotch is responsible for the peak value of the pressure fluctuation. This study illustrates that the geometry of R has a key impact on the unsteady shock interactions and, therefore, should be considered critically in practical applications, such as the cowl lip of a hypersonic inward-turning inlet.

충격파 풍동에서의 연속적 자유낙하 실험에 대한 적응적 실험 계획법 적용 연구

충격파 풍동에서의 연속적 자유낙하 실험에 대한 적응적 실험 계획법 적용 연구

The High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center (DLR)

The High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) is one of the major European hypersonic test facilities. It was commissioned for use in 1991 and was utilized since then extensively in a large number of national and international space and hypersonic flight projects. Originally, the facility was designed for the investigation of the influence of high temperature effects such as chemical and thermal relaxation on the aerothermodynamics of entry or re-entry space vehicles. Over the last years its range of operating conditions was subsequently extended. In this framework the main emphasis was to generate test conditions which allow investigating the flow past hypersonic flight configuration from low altitude Mach 6 up to Mach 10 in approximately 33 km altitude. The studies performed in HEG focused on the external as well as internal aerodynamics including combustion of hydrogen in supersonic combustion and the investigation of transition from laminar to turbulent hypersonic flow.

Convective heat-transfer rate and surface pressure distribution over a cone model at hypersonic speeds

In this article, experiments are carried out in a hypersonic shock tunnel with helium as driver gas and air as the test gas to obtain the convective heating rate and surface pressure distribution on a cone model placed at hypersonic speed. Test is performed in hypersonic shock tunnel for a flow Mach number of 6.5 at two different angles of attack, 0° and 5°. The sputtered thin film platinum sensors are used to measure the heat flux on a cone model. The measured heat-transfer rate compares well with theoretically estimated values using reference enthalpy method and computational fluid dynamics (CFD) simulation. The measured surface pressure compares well with CFD.

Aerothermodynamic effects of controlled heat release within the hypersonic shock layer around a large angle blunt cone

Effects of controlled heat addition into the high temperature, chemically reacting shock layer of a large angle (60°) blunt cone with a spherical nose have been experimentally investigated. The exothermic oxidation of ablated chromium from the surface of the cone at hypersonic Mach numbers triggers the heat release process. A conical skirt with a base diameter of 70 mm and a semi-apex angle of 30°, culminating into a nose of radius 30 mm, has been used as the test model. An in-house hypersonic free-piston driven shock tunnel facility, HST3, was used for the experiments at a stagnation enthalpy of 6.31 MJ/kg and a freestream Mach number of 9.84. The temperature distribution in the shock layer was experimentally measured by the two-color ratio pyrometry technique, using a Digital Single Lens Reflex (DSLR) camera as a pyrometer. The temperature field was corrected for discrete radiative line emissions obtained through emission spectroscopy. Surface heat flux measurements were taken using carefully calibrated thin film platinum heat transfer gauges mounted on an insulating Macor substrate. Shock stand-off distances were measured through Schlieren imaging of the flow using a high-speed camera at 20 000 frames per second, and also by a new intensity-scan based method using the processed color image from the DSLR camera. Calculations showed a 173 K rise in the temperature of the gas layer in the stagnation region due to chromium oxidation. The net surface heat flux on the blunt cone was also found to increase by about 31 W/cm2. The shock stand-off distance, as ascertained from Schlieren images, increased from about 3.82 mm (±1.4%) to 4.45 mm (±1.5%), a 17% rise. Analytical calculations, taking chromium oxidation reaction kinetics into consideration, to relate the total exothermic heat release to its distribution into various processes demonstrated that chromium oxidation releases about 78 W/cm2 energy in the stagnation region of the shock layer. 1.9% of this energy increases the temperature of the gas layer, 40% is convected back into the airframe, 0.4% is lost from the rear of the Macor substrate by conduction, and about 8% is radiated into the model. The remaining 50% of the heat was used up in pushing the shock layer away from the body by raising its density, thereby increasing shock stand-off distance.

Experimental research on high energy initiation technology for detonation driver

A shock tunnel with gaseous detonation driver is fit for simulating test gas with high total temperature and pressure, since the sound speed of the driving gas is comparatively high. Decreasing the sound speed of the driving gas can lead to initial detonation difficulties and is therefore deficient in low total temperature and high total pressure gas simulations. In this study, a new high-energy detonation initiation method, the closed-type ignition tube, is proposed, in which initial detonation is achieved in high-concentration nitrogen-diluted hydrogen-oxygen mixed gas and a low-sound speed driving gas is successfully obtained. By experiment, the dominant factors of igniter tube initiation energy are defined, and the design principles of closed igniter tube are given. Through experiments and calculation, the influence mechanism of the new method on the flow process of shock tubes/tunnels is clarified, and the principle of using the new initiation tube is identified. Using this new method, low total temperature and high total pressure test gas could be obtained, which provides a feasible method for expanding the ground test capability of hypersonic flight conditions.

Laminar–turbulent transition reversal on blunt ogive body of revolution at hypersonic speeds

The influence of flow parameters and nose radius on the location of laminar–turbulent transition is investigated. The model is an ogive-conical body of revolution having half angle about 9°. Experiments were conducted in a shock tunnel at Mach number 5. The transition location was diagnosed by the heat transfer rate distribution determined with the aid of luminescent temperature converters. It is shown that transition reversal can occur either (a) in the absence of turbulent wedges or (b) at a constant level of freestream disturbances. Both increasing and decreasing branches of Re∞,t (Re∞,R) dependency were observed at constant nose radius while varying only the unit Reynolds number.

A novel wavefront measuring camera for quantitative measurement of density in high-speed gas flows.

This paper presents the development of a novel wavefront measuring camera capable of detecting both the amplitude and phase of the captured light wave simultaneously. The main objective of the present work is to develop a simple "aim and shoot" camera system for quantitative estimation of density variations in high-speed gas flow fields. The interrogating beam which is a plane wave used here gets distorted by flow induced change in refractive index gradients. Wavefront distortion is quantitatively measured by inspecting the projected pattern through the embedded mask of a modified CMOS image sensor, which samples the incoming wavefront space continuously. Post-processing of the captured images through Fourier- and windowed Fourier transform schemes reveals the change in phase and amplitude of the captured wave. The captured phase of the wavefront is used in an iterative tomography scheme to estimate the density distribution of the flow field. The utility of the developed camera is demonstrated in the quantitative visualization of the high-speed flow fields around test objects subjected to hypersonic flows at Mach numbers 8.89 and 5.82 in hypersonic shock tunnel facility (HST2) and also to visualize the flow field generated at the exit of a convergent-divergent nozzle (Mach number 2.9). It is observed that the recovered quantitative density values from the experiments match well with the results obtained through computational fluid dynamic simulations demonstrating the proficiency of the proposed wavefront measuring camera for high-speed flow diagnostics.

Supersonic Combustion of a Scramjet Engine Using Hydrogen Fuel in Shock Tunnel

Experiments were carried out on a scaled model of scramjet engine with gaseous hydrogen as a fuel in a 1 m combustion-driven shock tunnel. The simulated enthalpy was , at a freestream Mach number of 6.8 with varying stagnation pressures (110–190 bar), dynamic pressures (72–130 kPa), and equivalence ratio (0.24–1.16). Detailed studies were carried out to tailor the shock tunnel. Experiments were carried out with air and as the test gases with hydrogen () injection to bring out the difference between reactive and nonreactive cases. Pressure measurement along the centerline of the engine and heat flux (through temperature measurements) inside the combustor was carried. The occurrence of supersonic combustion was ascertained through pressure measurements where no upstream influence was seen ahead of the strut injector and the same was corroborated by heat flux measurements. The pressure rise due to supersonic combustion downstream of the fuel injection was nearly three times as compared with nonreactive case. For the current case, when the combustor inlet pressure () exceeds 1 bar, a sharp rise in the pressure signature was observed downstream of fuel injection, whereas no such marked trace was seen when is below 1 bar.

Ethylene flame-holding in double ramp flows

In this work, ethylene flame-holding in supersonic flows was investigated using a shock tunnel. The experiments were conducted at flow stagnation temperatures ranging from 1270 to 1810 K. The two-dimensional test model consisted of a double-ramp inlet and a constant-area combustor with a recessed wall cavity. The two fuel injectors were located at the inlet and the other upstream of the cavity. Shadowgraph and flame chemiluminescence images were captured for optical visualization. The inlet injection images showed various flame-holding patterns. At 1270 K, the flame was not maintained. At 1540 K, the flame was maintained inside the cavity, and the condition provided continuous combustion during the steady flow. At 1810 K, strong flame signals were observed from the inlet to the cavity and downstream. At 1540 K, the inlet injection with a low fuel pressure showed a gradual flame quenching in the cavity during flow establishment. On the other hand, the same injection in the combustor showed flame-holding in the shear layer above the cavity. The results showed that the flame patterns are strongly influenced by the flow stagnation temperature and the location of fuel-injection.

Hand-operated shock tube and hypersonic shock tunnel

This manuscript outlines the invention of a miniaturized version of the conventional shock tube and hypersonic shock tunnel that is small enough to be mounted on a table top The shock tube named as Reddy tube works solely with the effect of human force and the driver gas is pressurized by manual action of a plunger rod Test can be conducted within a shock Mach number range of ndash The extension of the Reddy tube into a table top single pulsed hypersonic shock tunnel named as Reddy tunnel proves to be an important input into the field as it can be established as a fraction of the cost of a full fledged facility thereby providing access to such experiments to virtually any person in the experimental aerodynamic community

Velocity and NO-Lifetime Measurements in an Unseeded Hypersonic Air Flow

A laser-induced fluorescence (LIF)-based nitric-oxide flow-tagging technique was applied to measure both velocity and NO lifetime in a hypersonic shock tunnel from two experimental test runs. The results were supported by an analytical profile proposed in this paper that provides a way to correct velocity measurements under unknown systematic error sources. This procedure provided velocities with discrepancies lower than 3% for a total of five measurements, and lower than 2% when compared with that obtained from a linear fit. Additionally, the comparison between the proposed and experimental profiles allowed us to obtain the fluorescence NO lifetime from only one image.

Performance of Copper Calorimeter for Heat Transfer Measurement in High Enthalpy Shock Tunnel

Copper calorimeter, based on a calorimetric principle, offers a solution for heat transfer measurement in high enthalpy situation, especially in the erosive flow of high enthalpy shock tunnels. In this study, we numerically investigated the measuring performance of copper calorimeters. Non-ideal effects, such as heat loss to the insulator around and replacement of the average temperature of the copper element by the junction temperature, were discussed in detail. The influences of copper element thickness, copper/constantan wires thickness and sensor diameter were also estimated, with the aim to provide theoretical guidance for the design of copper calorimeter. In addition, corresponding experiments in JF10 high enthalpy shock tunnel were carried out against the data of coaxial thermocouples for verification. Results showed that the non-ideal thermal environment of a copper calorimeter (heat exchange with its surroundings) would result in a smaller measuring heat flux comparing to the one actually loaded; proper thickness of copper element matching the effective test time of shock tunnel was suggested. Besides, preliminary experimental results with corrections showed reasonable agreement with the heat flux of thermocouples, with an average deviation of 8%. Over all, this gauge developed extends and supplements the high enthalpy shock tunnel heat transfer measurements made by other techniques.