Impulse Facilities Research Articles

Nozzle Flow Characterization of the SNU Hypersonic Shock Tunnel

AbstractThis paper presents the operational capabilities and the characteristics of the nozzle flow of the Seoul National University Hypersonic Shock Tunnel (SHyST), a recently added impulse facility designed to produce high-enthalpy flows reaching up to 5 MJ/kg through hypersonic contoured nozzles. Within this investigation, emphasis is placed on the design and utilization of a pitot rake, serving as an essential instrument for quantifying pitot pressure and Mach number distributions throughout the test section. The experimental results confirm the presence of axis-symmetric behavior and spatial–temporal uniformity of the freestream Mach number at the nozzle outlet and along the test section. In addition, the study demonstrates a good agreement with the predictions of numerical simulation. Mach number validation is further supported by the measurement of shock standoff distance through schlieren visualization.

Investigation of Burst Pressure of Flat-Scored Metal Diaphragms in Hypersonic Impulse Facilities

Flat-scored metal diaphragms are essentially used in various hypersonic impulse facilities as quick-opening valves. Their burst pressure is a key parameter to optimize the performance of shock wave experimental devices and ensure the activation of overpressure safety devices. However, the conventional method to predict the burst pressure relies on time-consuming experiments that pose significant challenges for ultrahigh driving conditions. In this study, the finite element method (FEM) based on the Johnson–Cook model is adopted to predict the burst pressure of diaphragms used in a shock tube. The influences of the diaphragm thickness and groove depth on the burst pressure are analyzed. A simplified approximation based on the simulation results is obtained to estimate the burst pressure under a static load rapidly. This method is more generalizable than the existing equation and produces results in good agreement with experimental results. Furthermore, the burst pressure is investigated under different dynamic loads using the proposed FEM method. The results show that the dynamic load results in larger burst pressures than the static load, indicating that the burst pressure depends on the load type, the loading rate, and the magnitude of the applied forces.

Laser Absorption Sensor Targeting Potassium for Hypersonic Velocity, Temperature, and Enthalpy Measurements

A laser absorption-based sensor for hypersonic gas flows was developed, targeting the spectroscopic transition of atomic potassium near 770 nm. The sensor applies rapid-scanning tunable diode laser absorption spectroscopy to measure velocity from the Doppler shift and to infer temperature from the hyperfine-split transition lineshape. This sensor measured velocities and temperatures across three distinct conditions and six shots in the Hypervelocity Expansion Tube at the California Institute of Technology. Velocity and temperature were sampled at intervals, and temperature measurements were validated with a supplementary laser absorption-based sensor targeting carbon dioxide transitions near . Measured velocities across the three conditions ranged from 3.3 to 4.4 km/s, and measured temperatures ranged from 900 to 1600 K. The combined measurements were used to infer the freestream specific total enthalpy, which ranged from 7 to 10 MJ/kg. Because atomic potassium naturally forms in the test gas of many hypersonic impulse facilities, similar sensors may be widely applicable to facility characterization.

Flow Characterization of the UTSA Hypersonic Ludwieg Tube

The characterization of a hypersonic impulse facility is performed using a variety of methods including Pitot probe scans, particle image velocimetry, and schlieren imaging to verify properties such as the velocity, Mach number, wall boundary layer thickness, and freestream turbulence intensity levels. The experimental results are compared to the numerical simulations of the facility performed with Ansys Fluent to compare the design and operational conditions. The presentation of results in this manuscript is prefaced by a description of the facility and its capabilities. The UTSA Ludwieg tube facility can produce a hypersonic freestream flow with a Mach number of 7.2 ± 0.2 and unit Reynolds numbers of up to 200 × 106 m−1. The Pitot probe profiles of the 203-mm-square test section indicate a 152 ± 10 mm square freestream core with turbulence intensity values ranging from 1% to 2%. Schlieren imaging of the oblique shockwaves on a 15° wedge model provided an alternate means of verifying the Mach number. Particle image velocimetry and previous molecular tagging velocimetry results showed a good agreement with the Pitot probe data and numerical simulations in the key parameters including freestream velocity, wall boundary layer velocity profiles, and wall boundary layer thickness.

Direct measurement of skin-friction in shock/boundary-layer interaction flows

The direct measurement of skin friction at high speed is extraordinarily complicated due to the interference from various extraneous forces and the short test-time of high-speed facilities. The present study deals with the design and performance assessment of a skin friction sensor in a hypersonic flowfield with shock-boundary layer interaction, which is a predominant high-speed flow phenomenon. The experiment was conducted in the IIT Bombay Shock Tunnel (IITB-ST); the sensor was exposed to a flow with freestream Mach number, total enthalpy, and Reynolds number of 8.4, 0.69 MJ/kg, and 2.60 × 106 m−1, respectively. The shock-boundary layer interaction was induced by an oblique shock impinged on a flat plate, with a flow turn angle of 15°. The low magnitude average skin friction could be measured with an uncertainty of ±21%. The comprehensive study conducted confirmed the suitability of the skin friction sensor for high-speed applications in impulse facilities.

Characterization of reflected shock tunnel air conditions using a simple method

A new method to characterize air test conditions in hypersonic impulse facilities is introduced. It is a hybrid experimental–computational rebuilding method that uses the Fay–Riddell correlation with corrections based on thermochemical nonequilibrium computational fluid dynamic results. Its benefits include simplicity and time-resolution, and using this method, a unique characterization can be made for each individual experimental run. Simplicity is achieved by avoiding the use of any optical techniques and overly expensive numerical computations while still maintaining accuracy. Without making any assumptions to relate the reservoir conditions to the nozzle exit conditions, the work done characterizing four test conditions in a reflected shock tunnel is presented. In this type of facility, shock compression is used to produce an appropriate reservoir, which is then expanded through a nozzle to produce hypersonic flow. Particular focus is given to the nozzle exit total enthalpy where a comparison is made with the reservoir enthalpy obtained using the measured shock speed and pressure in the shock tube. Good agreement is observed in all cases providing validation of the new approach. Additionally, static pressure measurements showed clearly that conditions III and IV have a thermochemical state which likely froze shortly after the nozzle throat. Also, the nozzle flow is shown to be almost isentropic. Due to the simplicity of the current method, it can be easily implemented in existing facilities to provide an additional independent estimate alongside existing results.

Analysis of Reentry and Break-Up Forces from Impulse Facility Experiments and Numerical Rebuilding

This paper shows new findings for the break-up of large spacecraft during reentry into Earth’s atmosphere. The break-up scenario at high altitude drives the ground impact area for parts surviving the reentry. Therefore, based on a combined experimental and numerical analysis of the reentry of the International Space Station (ISS), loads at the module connections have been assessed. In this study, experiments were conducted in an impulse facility to determine the aerothermodynamic forces that apply to models of three components from the ISS. This paper marks the first approach to the experimental testing of complex geometric models coupled with a ground-to-flight scaling of the resulting internal stress. A high-speed schlieren setup was used to record the movement of the multibody free-flight models. A 3D positional analysis tracked the bodies separately. The experimental situation was simulated using the eilmer4 computational fluid dynamics code. A comparison of the experimental and simulated flowfield shows a reasonable agreement in shock structure and resulting forces. The forces were then scaled from testing to flight, extrapolated along a trajectory path, and transformed into spacecraft internal stress. While the resulting forces are significant, it is shown that these forces are significantly smaller than the yield strength of structural materials. Although no aeroheating was simulated in this study, the separation of the large segments of the ISS might not be driven by the mechanical forces alone. This study shows that comparatively simple shock tunnel experiments offer a comprehensive high-fidelity analysis of the interbody forces of complex multibody structures.

Formation of Bow Shock Waves Around a Blunt Body in Magnetohydrodynamics Flows

The diffraction of a shock wave over a stationary body is a problem of interest associated with the starting of shock tubes and expansion tubes, which are well suited to studies of hypersonic magnetohydrodynamic flows. However, these facilities are characterized by very short test times. The transient parameters during the establishment of the detached bow shock in such impulsive facilities are important for both data processing and experimental design. In the present study, based on the low magnetic Reynolds number assumption and dipole magnetic field distribution, the influence of magnetic field on the diffraction of an incident shock wave over a sphere was studied numerically. The incident shock Mach number ranges from 11 to 15 under different magnetic field intensities. Time histories of the shock-detached distance and stagnation pressure were first obtained. Moreover, the time needed to establish the steady flows over the sphere was also displaced against the magnetic interaction parameter. The larger the magnetic interaction parameter, the longer the time needed to establish a steady bow shock for experiments, which further challenges the facilities to have sufficient test times for conducting hypersonic magnetohydrodynamic flow experiments.

Plasma preheating technology for replicating planetary re-entry surface temperatures

AbstractArc-jet facilities have been the norm for ablation experiments used to calibrate computational models to date. However, the arc jet has a few major limitations and challenges, including non-uniform enthalpy distribution, non-equilibrium state, change of surface quality during testing and the extent of oxidation, to name but a few. A novel plasma technique for preheating axisymmetric heatshield samples in hypersonic impulse facilities is presented herein. The major aim of this innovative work is to help reduce the large variations of ablation rate predictions, space vehicle materials and missile design/testing, obtain strongly coupled hypersonic boundary layers and achieve lower cost of aerothermodynamics experiments. This present work remains one of the most highly anticipated solutions to maximise payload success and replicate high surface temperatures identical to those experienced by real flight vehicles. This work makes a useful contribution to re-entry studies under conditions that replicate the characteristics of re-entry flights. Future applications for the technique are expected to be found in hypersonic impulse facilities that can simulate the true flow energy under re-entry conditions.

Measurements of Reflected Shock Tunnel Freestream Nitric Oxide Temperatures and Partial Pressure

This paper reports on measurements of freestream nitric oxide (NO) rotational and vibrational temperatures and partial pressures, collected in the Caltech T5 reflected shock tunnel. Quantum cascade lasers, emitting mid-infrared light resonant with fundamental rovibrational NO transitions, were directed through the supersonic (Mach ) freestream flow. Tunable diode laser absorption spectroscopy (TDLAS) was used to measure the path-averaged rotational and vibrational temperatures of NO in the flow, in addition to the NO partial pressure. The temperature measurements demonstrate strong evidence of NO rotational and vibrational equilibrium during the 1 ms test period. Agreement between vibrational and rotational temperatures was observed in all experiments, including one and four experiments, during and after the nominal test time. Absorption from CO and was also observed in the TDLAS measurements, though their concentrations cannot be accurately estimated. The goal of these and future experiments is to develop and demonstrate TDLAS experimental strategies for high-enthalpy impulse facilities and to help to inform improvements of existing models and solvers used for prediction of freestream conditions.

A skin-friction balance for shock tunnel applications

An acceleration-compensated skin-friction balance based on the floating-head technique was developed to measure skin friction in a hypersonic impulse facility. The balance consisted of measurement and acceleration units, the outputs of which can be coupled using a full-Wheatstone bridge to obtain pure shear output. The working of the balance was tested on a flat plate model exposed to a laminar flow with Mach number, stagnation enthalpy and Reynolds number of 8.2, 0.65 MJ kg−1 and 2.9 × 106 m−1, respectively. The measured wall shear stress agreed reasonably well with that predicted by Reynolds analogy.

Technique Development for Investigating Axisymmetric Ablation Models in Hypersonic Impulse Facilities

Arcjet facilities have been used extensively for ablation and materials characterization but arcjets do not correctly simulate the hypervelocity flow during reentry. Impulse facilities can simulate aspects of hypervelocity reentry flows, but cannot duplicate the extended-duration thermal environments. To address such deficiencies in short-duration testing environments, a novel plasma technique for preheating axisymmetric samples of heat-shield materials has been developed. The heating technique has been demonstrated using a short-duration blow-down arrangement at the University of Southern Queensland. A 50-mm-diameter graphite disk sample with a uniform thickness of 2 mm was heated from the downstream side with plasma to approximately 2500 K, and then it was exposed to a cold Mach 4.5 flow. The disk was mounted on an experimental probe that was very similar to the European standard probe, normal to the flow. Future applications for the technique are expected to be found in expansion tube facilities that can simulate the true flow energy under reentry conditions.

Pulse-burst spontaneous Raman thermometry of unsteady wave phenomena in a shock tube.

A high-speed temperature diagnostic based on spontaneous Raman scattering (SRS) was demonstrated using a pulse-burst laser. The technique was first benchmarked in near-adiabatic ${{\rm H}_2} \text{-} {\rm air}$ flames at a data-acquisition rate of 5 kHz using an integrated pulse energy of 1.0 J per realization. Both the measurement precision and accuracy in the flame were within 3% of adiabatic predictions. This technique was then evaluated in a challenging free-piston shock tube environment operated at a shock Mach number of 3.5. SRS thermometry resolved the temperature in post-incident and post-reflected shock flows at a repetition rate of 3 kHz and clearly showed cooling associated with driver expansion waves. Collectively, this Letter represents a major advancement for SRS in impulsive facilities, which had previously been limited to steady state regions or single-shot acquisition.

Evolution of heat transfer at the stagnation point during the detached bow shock establishment

The diffraction of a shock wave over a stationary body is a problem of interest associated with the starting of shock tubes and expansion tubes which are well suited to studies of hypersonic flows. However, these facilities are characterized by very short test times. The transient parameters during the establishment of the detached bow shock in such impulsive facilities are important for both data processing and experimental design. In the present study, numerical simulations are conducted to investigate the diffraction of a normal shock wave over a sphere and the subsequent transient phenomena in a viscous perfect-gas flow field. The incident shock Mach number ranges from 3 to 5 with a specific heat ratio of 1.4. Based on the theoretical description of the reflected shock position during bow shock formation, approximate solutions for the time histories of the stagnation-point heat flux are also derived. The analytical and numerical results agree well. The results show that the stagnation-point pressure and heat flux approach their steady-state values much more rapidly than the shock detachment distance does.

In situ nozzle reservoir thermometry by laser-induced grating spectroscopy in the HELM free-piston reflected shock tunnel

Experimental determination of test gas caloric quantities in high-enthalpy ground testing is impeded by excessive pressure and temperature levels as well as minimum test timescales of short-duration facilities. Yet, accurate knowledge of test gas conditions and stagnation enthalpy prior to nozzle expansion is crucial for a valid comparison of experimental data with numerical results. To contribute to a more accurate quantification of nozzle inlet conditions, an experimental study on non-intrusive in situ measurements of the post-reflected shock wave stagnation temperature in a large-scale free-piston reflected shock tunnel is carried out. A series of 20 single-shot temperature measurements by resonant homodyne laser-induced grating spectroscopy (LIGS) is presented for three low-/medium-enthalpy conditions (1.2–2.1 MJ/kg) at stagnation temperatures 1100–1900 K behind the reflected shock wave. Prior limiting factors resulting from impulse facility recoil and restricted optical access to the high-pressure nozzle reservoir are solved, and advancement of the optical set-up is detailed. Measurements in air agree with theoretical calculations to within 1–15%, by trend reflecting greater temperatures than full thermo-chemical equilibrium and lesser temperatures than predicted by ideal gas shock jump relations. For stagnation pressures in the range 9–22 MPa, limited influence due to finite-rate vibrational excitation is conceivable. LIGS is demonstrated to facilitate in situ measurements of stagnation temperature within full-range ground test facilities by superior robustness under high-pressure conditions and to be a useful complement of established optical diagnostics for hypersonic flows.

Numerical modelling of thin film gauge for short duration thermal measurement

Abstract The numerical study of direct heat conduction through thin film heat flux gauge is carried out to see how closely it resembles with short duration experimental thermal data. The thermal data obtained from transient impulse facilities at Von Karman Institute, Belgium (case1) and Oxford University (case2) is taken from literature as a reference for validation purpose. The numerically simulated temperature data is compared with experimental data. A good agreement is observed. Further, heat flux is recovered from temperature data using the cubic spline technique for comparative study. The peak deviation from experimental heat flux data for case 1 and 2 is found to be 6% and 7.2% respectively. All these deviations are in the acceptable range of 8%. This deviation is also accounted by the fact that direct heat condition modelling is used instead of numerical modelling of impulse facilities. Thus, direct heat conduction modelling is acceptable approach to predict the transient thermal data with admissible errors.

Intelligent Force-Measurement System Use in Shock Tunnel.

The inertial vibration of the force measurement system (FMS) has a large influence on the force measuring result of aircraft, especially on some tests carried out in high-enthalpy impulse facilities, such as in a shock tunnel. When force tests are conducted in a shock tunnel, the low-frequency vibrations of the FMS and its motion cannot be addressed through digital filtering because of the inertial forces, which are caused by the impact flow during the starting process of the shock tunnel. Therefore, this paper focuses on the dynamic characteristics of the performance of the FMS. A new method—i.e., deep-learning-based single-vector dynamic self-calibration (DL-based SV-DSC) of an impulse FMS, is proposed to increase the accuracy of aerodynamic force measurements in a shock tunnel. A deep-learning technique is used to train the dynamic model of the FMS in this study. Convolutional neural networks with a simple structure are applied to describe the dynamic modeling so that the low-frequency vibration signals are eliminated from the test results of the shock tunnel. By validation of the force test results measured in a shock tunnel, the current trained model can realize intelligent processing of the balance signals of the FMS. Based on this new method of dynamic calibration, the reliability and accuracy of force data processing are well verified.

Experiments on water vapour condensation within supersonic nozzle flow generated by an impulse tunnel

An impulse facility for analysis of water vapour nozzle flows using both shock tunnel and Ludwieg tube operating modes has been developed and tested at the University of Southern Queensland. Unique high-speed flow visualisation of the water vapour condensation shock has been acquired in the throat region of a nominally two-dimensional convergent-divergent nozzle with a 120 mm2 throat area. This paper presents the facility performance and the time-resolved visualisation results for the nozzle flow, and the results are analysed with the aid of image post-processing tools and quasi-one-dimensional (Q1D) thermodynamic calculations. The experiments have produced qualitative and quantitative data on the flow conditions at four different relative humidity values from 25% to 100%, and over a range of nozzle supply temperatures between 293 K and 343 K. The direct measurement of the pressures and wave speeds within the tunnel resulted in a good agreement with the ideal gas Q1D calculations, particularly for the first incident shocks and expansion waves. Differences progressively increased with the number of reflections, due to the effect of viscous dissipation and the non-ideal end-wall interaction caused by the presence of the nozzle inlet. Moreover, flow visualizations of the location and orientation of both the weak disturbance waves and the condensation shock enabled the investigation of the supersonic water vapour condensation within the nozzle of this impulse facility. Based on these measurements, the nozzle flow downstream of the condensation shock was demonstrated to have a lower Mach number relative to the dry nitrogen flow case. This phenomenon is due to the phase change heat release, which appeared to be more significant when the condensation shock intensity was higher. Although further development of the apparatus is possible, the impulse test-rig configuration that we have developed permits a relatively high flow rate of test gas for a short period of time, leading to significantly reduced operating expenses relative to a continuous flow facility with the same mass flow rate. It is therefore suggested that future studies further improve and test the impulse facility techniques for studying condensation phenomena as a potential alternative to high-power, continuous flow facilities.

Experimental validation of a test gas substitution for simulating non-equilibrium giant planet entry conditions in impulse facilities

Scaled giant planet entry testing is beyond the simulation capability of most impulse facilities, which limits the ground testing data available for computational model validation. Substituting the He in their predominately H/He atmospheres with Ne has been used in the previous experiments. This enables higher-temperature shock layers to be recreated using the same facility performance. For binary hydrogen dissociation and ionization reactions, the substitution gives similarity with non-equilibrium processes, but has never been experimentally validated by direct comparison with H/He relaxation data in a region where the facility simulation capabilities in using H/He and H/Ne overlap. This work demonstrates and validates a scaling method based on the substitution in terms of flow relaxation. A H/Ne condition was designed to recreate the post-shock relaxation of a H/He condition in our X2 expansion tube. The two conditions were experimentally tested with cylindrical test models, and shock layer radiation from the Balmer series was measured. The spectroscopic data using the H/Ne condition show a successful recreation of the non-equilibrium processes of the target H/He condition, experimentally validating the scaling method. High-speed images demonstrate that the associated radiation field is also correctly reproduced. With the use of this substitution, the facility can now be used to simulate high-speed giant planet entries with confidence, and the available ground testing envelope is extended to much higher speeds.

Magnetohydrodynamic drag force measurements in an expansion tunnel using a stress wave force balance

Magnetohydrodynamic (MHD) aerobraking has the potential to significantly reduce peak heat loads for high-speed planetary entry, due to increases in total drag force in the early phase of the trajectory. Establishing accurate techniques for measuring MHD drag force in ground testing facilities is a necessary first step for investigating this phenomenon. The objective of this paper is to demonstrate the methodology required for using a stress wave force balance (SWFB) to measure MHD drag force in an expansion tunnel. Whilst the SWFB is an established technique for measurement of short duration forces in impulse wind tunnel facilities, conventional designs exhibit unusually low signal-to-noise ratios in MHD ground testing flow fields, due to electromagnetic interference from the plasma in the shock layer. This paper investigates the effect that sting design, amplifier setup, and strain gauge location have on the signal-to-noise ratio of a SWFB. Experimental validation of each design was undertaken in the X2 expansion tunnel with a high enthalpy ( $$\sim 16\,\hbox {MJ}\,\hbox {kg}^{-1}$$ ) argon test flow by comparing the results to those obtained using an accelerometer-based force balance. The results demonstrate that maximum signal-to-noise ratio is obtained when: the sting is made from a low stiffness material such as polycarbonate to maximise strain; the charge amplifier is located near the strain gauge inside the model and is grounded at the data acquisition system to minimise electrical noise; the strain gauge is located close to the applied load to delay the return of reflected stress waves from the free end. MHD drag forces measured in this study varied up to approximately 3 N, and measurement uncertainty was found to be approximately $$\pm \, 0.2\,\hbox {N}$$ . Overall, successful adaptation of the stress wave force balance technique to MHD ground testing was achieved, and it has been shown to be a viable alternative to accelerometer-based techniques.