Expansion Tube Magnetohydrodynamic Experiments with Argon Test Gas

Expansion tubes perform an important role in hypersonic flight research by providing a ground-based testing facility in which scale models can be tested at simulated flight conditions. High enthalpy and high velocity flow conditions are generated via an unsteady expansion process which accelerates test gas to hypersonic conditions. Unfortunately, due to the high velocities and low densities created by this expansion process, test times are shortened and very thick boundary layers develop, reducing the useable core flow. This restricts the size of the models that can be tested, as well as the ability to obtain sufficient or accurate experimental data.Previous studies have shown that a nozzle could be added to an expansion tube to not only expand the core flow diameter but also increase the test time. This included the successful implementation of a nozzle on the University of Queensland's X2 expansion tube [31]. In this thesis, a full-capture, contoured, shock-free nozzle has been designed for the X3 expansion tube. The X3 nozzle contour was scaled from the X2 nozzle contour and was originally 3.0075m long.CFD modelling of the X3 expansion tube with the nozzle attached was completed using a hybrid method coupling a one-dimensional and axisymmetric simulation in order to predict the flow conditions produced by the nozzle. Code package L1d [16] was utilised for the onedimensional analysis and MB_CNS [14], for the axisymmetric analysis. The core flow diameter was increased from 150mm to approximately 250mm and the test time increased by 70% to 1ms; confirming that the addition of a nozzle to the X3 expansion tube will expand the core flow and increase the test time. The nozzle was also examined in relation to its potential to be truncated and maintain its core flow diameter and test time. It was determined that the nozzle could be truncated to a length of 2.5m with no observed loss to the core flow size and test time achieved.The X3 nozzle is under construction. A fibre-wound composite construction technique is being utilised. This construction method was previously used by Jacobs in the fabrication of a nozzle for the T4 reflected shock tunnel [18] at the University of Queensland, although, has not been used for expansion tube nozzles. This technique has proved effective in providing a low-weight, reduced cost alternative to metals.The mandrel for the fibre-winding process was fabricated from high-density polyurethane foam in order to overcome issues of weight and surface instability experienced with other materials. This material had not been previously used for this application and proved very successful in this providing a rigid and stable surface that was suitable for machining as well as winding over.On installation of the nozzle on the facility, the testing of larger, more accurate models is possible. Test times will also be longer than previously observed, enabling the collection of more data, further improving the capacity to effectively model hypersonic flight conditions with this facility. This will benefit future ground based hypersonic flight research to beconducted in the X3 expansion tube facility. Overall, the work completed in this thesis has furthered the development of the X3 expansion tube and provided a foundation for greatly expanding the capabilities of this facility.

A set of experiments was conducted with the aim of evaluating total heat flux mitigation using magnetohydrodynamic flow control over ionized shock layers surrounding planetary entry vehicles. The University of Queensland’s X2 expansion tube was used to generate a flow condition representative of a Mars return trajectory. Spherical permanent magnets were fitted into a scaled test model based on NASA’s Stardust capsule. Thermocouples located at the stagnation point and shoulder of the test model, as well as thin-film heat transfer gauges located in the aft section, measured total heat flux for both magnetic and nonmagnetic experiments. High-speed imaging of the shock layer showed a significant increase in shock standoff when using the magnet, and polynomial fits were derived, describing standoff distance along the model’s forebody. Despite significant noise affecting the heat flux gauges signals for the magnetic experiments, a simple methodology allowed to compare heat flux during the useful test time. It was found that the applied magnetic field significantly reduces total heat flux at the stagnation point and shoulder of the test model, although heat flux to the aft section was found to significantly increase. Reasonable agreement with the numerical predictions was found for the aft section of the test model, but discrepancies were observed at the other locations.

Future space missions, such as return from Mars, will require entry speeds of the order of . These harsh conditions are dominated by radiative heat flux and have large modeling uncertainties related to radiative coupling, such as precursor radiation and ablation radiation coupling. These conditions have also never been tested in a flight scenario and have limited ground testing data available. This study investigates the feasibility of generating these conditions in the X2 expansion tube at the University of Queensland so that these flows can be tested over scaled test models for the study of shock relaxation, heat flux, shock standoff, or ablation. This paper presents a theoretical performance map of the X2 facility for the generation of these conditions before it presents some of the unique issues that may be encountered when generating and investigating them. Some preliminary experimental data are then presented, showing that these 15 km/s equivalent flight speed test conditions can be generated in X2 within the current performance limits of the facility.

A major design challenge for re-entry capsules lies in the modelling of convective and radiative heat transfer to the surface of the vehicle. At certain points on superorbital re-entry trajectories, up to 40% of the total radiative heat flux is contributed by the vacuum ultra-violet (VUV) spectral range and it is in this spectral range that the largest uncertainties lie. The high level of uncertainty in the VUV is a result of a lack of published experimental data due to difficulties encountered in measuring radiation in the VUV, such as strong absorption by most optical materials and air. Additional complexities of the VUV spectral range include its strongly self-absorbing nature and spectral line broadening. The primary goal of this study was to obtain calibrated spectral measurements in the VUV that enable the investigation of physical processes occurring in the shock layer that influence the incident radiative heat flux. In particular, the issues to be investigated were the variation in spectral radiance observed across a shock layer compared to the spectral radiance measured through the surface, the effects of self-absorption on spectral line intensity and the broadening of spectral lines in the VUV as a function of depth of radiating flow field. The measurements made across and through the surface of a model provide the first set of calibrated experimental results for the validation of computational codes used to predict incident radiative heat flux. Measurements made with a varying depth of radiating flow field provide a unique set of experimental data for the validation of radiation transport models and broadening coefficients. This study also used computational simulations to investigate the accuracy of a flow field solver coupled with two reaction rate schemes and compared the spectra produced using Specair with experimentally measured values. To achieve these goals, an optical system was designed to measure the VUV radiative emission produced around a blunt two-dimensional model in a spatially resolved manner across the shock layer. Spatial resolution allowed for spectral measurements to be made in both the equilibrium and non-equilibrium parts of the shock layer. A second optical system was designed to obtain measurements of VUV radiation incident on the surface of the model. This system incorporated a window in the surface with a mirror housed within the model to deflect the radiation out of the test section and into the detection system. To effectively vary the depth of the radiating flow field, the length of a two-dimensional model was varied, changing the depth of the shock layer being observed. The X2 expansion tube was used to create the high enthalpy flows required to produce radiating shock layers. Two flow conditions were created for this study that represented flight equivalent velocities of 10.0 km/s and 12.2 km/s. The spectroscopy system utilized for this study consisted of an evacuated McPherson NOVA 225 spectrometer coupled to an Andor iStar VUV enhanced intensified charge coupled device. An evacuated light tube sealed with a magnesium fluoride window was required to extend the evacuated light path to the model and avoid any absorption by molecular oxygen. An in-situ calibration of the VUV spectroscopy system was conducted using a deuterium lamp located in the position of the radiating shock layer. The integrated incident spectral radiance measured through the surface of the model between 115 nm and 180 nm was 0.744 W/cm2 sr for the 10.0 km/s condition and 12.3 W/cm2sr for the faster 12.2km/s condition. [...]

This paper explores the operating envelope for the University of Queensland’s free-piston driven X2 expansion tube to determine if flow conditions can be produced that generate a significant magnetohydrodynamic interaction. Numerical calculations are presented to demonstrate the existence of viable operating conditions to study magnetohydrodynamics. Argon is selected as the test gas due to its simple chemistry and low ionization enthalpy. A candidate flow condition is selected for detailed analysis and experimental validation. Finite-rate reacting argon numerical simulations indicate that relatively short shock-layer length scales for the models tested in X2 limit the degree of argon ionization that is generated. However, a strong interaction remains available, and use of a larger facility can fully mitigate this issue. It is shown that expansion tubes are capable of generating the required flowfield for these experiments, and important considerations for design of suitable test flows are identified.

Availability of reflected shock tunnels capable of testing large-scale models and producing long-duration test flows is currently limited. The X3R free-piston reflected shock tunnel is in the final stages of development as an alternate operating mode of the pre-existing X3 expansion tube at The University of Queensland. A set of analytical and numerical procedures have been developed to aid in optimization of new operating conditions. This paper summarizes the condition development process and provides details of three candidate Mach 7 operating conditions theoretically capable of generating at least 10 ms of test time for full-scale scramjet engine testing. The facility’s atypical geometry, arising from its original design as an expansion tube, and the required compromises made during condition development are discussed.

The University of Queensland has completed extensive research into the flow of fluid around objects to simulate the re-entry of spacecraft into atmospheres and flow through scramjets (Jacobs et al., Unknown). This has been made possible through the use of expansion tubes which can reach gas flow velocities in excess of 10km/s (Jacobs et al., Unknown).The expansion tubes operate on gas expanding from an extremely high pressure through to a very low pressure. Objects are placed in the path of expansion to take measurements (Davey, 2006).Due to an X3 expansion tube upgrade at UQ the negative pressure volume (dump tank) is insufficient in size to keep sub-atmospheric pressures at a maximum operational envelope. This thesis looks at the design of a new dump tank.The new dump tank has been designed to have an optimised volume in order to minimise negative pressure pumping time yet still hold a sub-atmospheric pressure after a maximumenvelope shock has been created (R Morgan 2007, pers. Comm., March).Comparisons between measured pressure and predicted pressure were used on the X2 and T4 expansion tubes to find an accurate technique to calculate the required new dump tank volume. The most accurate analytical technique investigated was found to assume isentropic expansion of the gas with a specific heat ration of 1.4. One-dimensional modelling was run with unrealistic computation times and axisymmetric modelling was found to over estimate the spent gas pressure.

Designs for planetary reentry experiments using additive manufacturing technology have been investigated in X2 expansion tube at the University of Queensland. A 1:5 scale Hayabusa model with 12 thin film sensor mounts for instrumentation was proposed for faster tunnel turnarounds. Quasi-2D configurations were also designed with hub and shell ideas so that the shell can be discarded after testing and reproduced for faster re-fire rate. Additionally, 2D and 3D IXV models were SLA printed in resin and tested in X2 expansion tube. Due to different camera settings for both models, it was difficult to compare the flow of two 3D-printed models with steel baseline model. Both models survived the experimental test-time, however, all thin films attached to them were destroyed after testing. A future possible extension to this work is a `tree' 1:5 scale Hayabusa model using generative design tools in Fusion 360 to reduce mass of this component. Lastly, new material which has similar properties to Sample Return Capsule's ablative heat shield of NASA Stardust mission is Polyether ether ketone (PEEK) with chopped carbon fibre (CF). Current commercial 3D printing CF/PEEK is CarbonX product with 10% high-modulus chopped carbon fibre. In future, experimental studies of these new designs could be useful for this research.

Testing of the first heatshield of the Fire II reentry vehicle was performed in the X1 superorbital expansion tube at The University of Queensland. The test model was a 1:28 scale replica of the flight vehicle, and incorporated three thermocouples: stagnation and two radial. A trajectory point towards the end of the first experimental testing period, at a total flight time of 1639.5s, an altitude of 61.5Km and velocity 11.1km/s was simulated in the expansion tube. Stagnation point heat transfer was obtained using a fast response coaxial type E thermocouple. In the current analysis the convective and radiative heating components were treated independently, where the convective component was scaled with the length scale and the absolute value of the radiative heat transfer was held constant. From this, the overall contribution of the radiative heat transfer to the total heat rate is decreased in the expansion tubes from an 18% contribution in flight to less than 1%, whereas the convective component was increased by a factor of 28. This results in the convective heat transfer being the major mode of heat transfer in the experimental models. From the Fay and Riddell empirical convective heat transfer correlation it was shown that the parameter Ch√Re should remain constant between the flight and experimental tests provided ρL scaling is maintained. Results from the current study show good agreement with the convective heating component of the flight vehicle and the Ch√Re values are in agreement to within 20% of the flight results. The results obtained in this study give a strong indication that the relative radiative heat transfer contribution in the expansion tube tests is less than that in flight, supporting the analysis that the absolute value remains constant with ρL scaling.

The aerodynamics of entry into gas giants are dominated by thermochemical processes occurring in the shock layers. These processes are specific to the gas giants, with their unique atmospheric composition of hydrogen and helium, and are not well understood. This paper used the condition with the maximum enthalpy which can be achieved in the expansion tube, X2, as well as increased helium content in the test gas, to generate a radiation field representative of blunt body Saturn entry. These are the first spectrally resolved radiation measurement ever performed in an impulse facility to study simulated Saturn entry. Both high speed imaging and spectrometry were employed to measure the atomic hydrogen radiation emitted from Balmer series of hydrogen atoms. The bright region in the radiation field merely exists near the model surface, indicative of the slow ionization process,which causes the electron number densities to build up only near the surface. From the central streamline spectral data, H-α was the strongest transition, while other transitions (n < 3) were negligible after the shock, and became brighter near the body. Both the horizontal radiance profile and high speed image were employed to locate the focused slices in vertical shots. It was also found by the Specair calculation that self-absorption is more likely to occur in H-α radiance than in H-β radiance with the line of sight in this study.

The University of Queensland (UQ) is currently developing high Mach number, high total pressure scramjet flow conditions in its X2 and X3 expansion tube facilities. These conditions involve shock-processing a high-density air test gas followed by its unsteady expansion into a low-pressure acceleration tube. This relatively slow shock-processing requires the driver to supply high pressure gas for a significantly greater duration than normally required for superorbital flow conditions. One technique to extend the duration is to operate a tuned free-piston driver. For X2, this involves the use of a very light piston at high speeds so that, following diaphragm rupture, the piston displacement substitutes for vented driver gas, thus maintaining driver pressure much longer. However, this presents challenges in terms of higher piston loading and also safely stopping the piston. This article discusses the tuned driver concept, the design of a very lightweight but highly stressed piston, and details the successful development of a new set of tuned free-piston driver conditions for X2.

The experimental activities at Shock Wave Research Center of Tohoku University using a ballistic range are reviewed. First, the shock stand-off distance for a sphere and the shock shape over a sharp cone are measured in the hypersonic speed regime. These measurements are made for 30 mm diameter models, flight speeds between 2.5 and 4.0 km/sec and ambient pressure 4.2 and 76 mmHg. The flow field is visualized by the schlieren method using a short-pulse Nd-YAG laser as the light source. Shock stand-off distance and shock shape are shown to vary with the ambient pressure, demonstrating the effect of the chemical nonequilibrium. Secondly, hypervelocity impact tests are conducted for simulation of space debris impacts against a bumper shield. The bumpers are made of metal and nonmetal materials. The projectile made of high density polyethylene has a cylindrical shape of 14 mm diameter and 3.7 g in weight. The impact velocity is about 4 km/sec. The debris cloud is visualized by shadowgraph. The nonmetal bumper shields used in this study are more effective than the metal bumper in regard to both the bumper mass and defensive capacity. Introduction The ballistic range is one of the readily available impulse facilities for the fundamental investigation of the hypersonic flow phenomena. The so-called two stage light gas gun can accelerate a model to hypersonic speeds. The facility is used mainly for the study of hypersonic aerothermodynamics and for the hypervelocity impact tests. In Shock Wave Research Center of the Institute of Fluid Science of Tohoku University, experiments are in progress in a ballistic range with a 14 mmand 30 mm-bore two stage light gas gun. First, the shock stand-off distance for sphere and the shock shape over sharp cone are measured in the thermochemical nonequilibrium regime. In the past, a large number of experimental investigations on high temperature real-gas effects have been carried out in the shock tunnel. However, there exist uncertainties in the characteristics of the flow produced in a shock * Graduate student f Professor, Institute of Fluid Science, Tohoku University; Member AIAA. Copyright ©1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. tunnel in real-gas regime, and therefore it is necessary to validate the facility.Theoretically, the real-gas effects are being studied using computational fluid dynamics (CFD). However, the CFD calculation of nonequilibrium flow encounters many difficulties because of the complex processes of chemical reactions in the nonequilibrium state, particularly in the case of air. Therefore reliable experimental data are needed for validating shock tunnel results and CFD. One of the most appropriate reference quantity is the shock stand-off distance for a sphere and shock shapes for a sharp cone, because the heat absorption by the real-gas phenomena behind shock wave sensitively affects the shock stand-off distance and shock shapes.'' Ballistic ranges are the only impulse facility capable of producing a reliable data, because, in a ballistic range, a model flies in a quiescent atmosphere. Measurement of shock stand-off distance for sphere has been made by Lobb in the ballistic range (see Figure I). His experiment misses the range of the binary scaling parameter pR above 2X10 kg/m and the flight velocity below 4 km/sec. The accuracy of Lobb's experiment is unclear because of the relatively long pulse duration of the light source used for flow visualization and small size of the models. Therefore, we conduct an experiment using a Nd-YAG laser of very short pulse duration for the light source and use 30 mm-diameter models. The shock shapes have been also measured in real gas regime. However, these measurements are conducted in high enthalpy shock tunnels and expansion tubes.' As mentioned above, the data

The importance of vacuum ultraviolet (VUV) aerothermodynamic heating during reentry is presented and the uncertainties involved in this region of the spectrum and their consequent result on heat shield design is discussed. A method developed at The Centre for Hypersonics capable of VUV emission spectra measurements from across and through the surface of a blunt model is outlined. An in situ calibration is applied to the data measured using this system to achieve fully calibrated spectral data at two superorbital conditions representing high-speed points on a re-entry trajectory. To date, this is the first study to present calibrated through surface VUV spectral data at high speed superorbital trajectory points. In comparing through surface and across surface data, it is possible to calculate a correlation between the VUV radiation observed across a shock layer, as measured in shock tunnels, and the radiation incident on the surface. This correlation will enable shock tunnel data to be more accurately applied in the design of spacecraft heat shields and provide insight into VUV radiation heat transfer through boundary layers.

Experiments were conducted using the X2 expansion tube at two conditions representative of flight velocities of 10.0 and for a vehicle entry into Earth’s atmosphere. Calibrated spectral measurements were made of radiating shock layers across and through the surface of a model. These measurements provided a unique dataset for the validation of computational codes used in the calculation of incident vacuum ultraviolet radiative heat flux. Computational simulations at equilibrium gas chemistry were conducted using the Specair program and showed good agreement with the measurements.

During reentry into the atmosphere, a vehicle can encounter extreme convective and possibly radiative heating loads. Such spacecraft must rely on a carefully designed thermal protection system to maintain the integrity of the underlying structure, thus ensuring the survival of any cargo or crew it may contain. Despite many past developments and research efforts, to this day there remain large uncertainties associated with the prediction of heat shield performance. In particular, accurate modelling of ablation phenomena is critical to minimising uncertainties, reducing excess weight, and increasing safety. This includes thermochemical ablation by oxidation, nitridation, and sublimation, as well as the mechanical removal of material via spallation. Due to the high costs associated with in-flight testing, and the difficulty of mounting precision instruments and advanced diagnostics on flight vehicles, ground-based experiments stand as a cost-effective method for reducing modelling uncertainties. Ablation testing is generally conducted in long-duration facilities such as arc jets, which are well suited to investigating in-depth material response due to their ability to provide representative heating rates for extended periods. They are, however, unable to reproduce a realistic in-flight shock layer due to low Reynolds numbers and, depending on the facility type, a high degree of thermal and chemical nonequilibrium in the free-stream. In contrast, impulse facilities such as shock tubes and their variants are well suited to reproducing realistic flight conditions with free-streams that are a closer match for flows dominated by binary processes in terms of Reynolds number, temperature, and chemical composition. Due to their extremely short test times, however, ranging from the order of 10 us to several milliseconds, they are unable to aerothermally heat test models up to representative in-flight temperatures, or recreate the quasi-steady heat and mass flux balance of flight. This key limitation of impulse facilities was recently overcome in experiments using the University of Queensland’s X2 expansion tunnel by electrically pre-heating samples to temperatures approaching 2500K immediately prior to a test. In this work, the pre-heating methodology was enhanced to generate maximum surface temperatures approaching 3300K in order to study sublimation phenomena. It was found that although sublimation-regime temperatures could be achieved, the relevant surface processes themselves were suppressed due to the high pitot pressures of the conditions used. Instead, the nitridation process was studied by observing air shock layer CN emissions with pure graphite models heated to surface temperatures ranging from 1770–3300K. These results were compared with numerical simulations which predicted a monotonic increase of emissions with surface temperature for all models considered, whereas the experiments showed an initial increase from 1770–2500K and then remained relatively constant from 2500–3300K. The simulations also suggested that a significant portion of the observed CN emissions were not a result of direct nitridation, but instead due to surface CO production and subsequent gas-phase interactions to form CN. A final series of experiments was conducted using a novel model design whereby the heated graphite strip was embedded into a steel base structure in order to relieve edge effects, with the aim of producing quasi two-dimensional flow along the strip. The degree of flow two-dimensionality was found to be questionable due to visible cross-flow and a shock stand-off 10% smaller than a nominal two-dimensional case, although it was much closer than a heated strip on its own. These tests were conducted in both air and nitrogen flows to further investigate the coupling effect of surface CO production on the levels of CN emissions via gas phase reactions, as suggested by the previous comparisons with numerical simulations. Based on the relative levels of CN emissions between the two test conditions, it was concluded that although CO interactions were a non-negligible contributor to CN production, the majority was in fact due to direct surface nitridation. In terms of overall trends, the nitrogen condition results showed continually increasing CN emissions with surface temperature. For air, however, they reduced with temperature above 2500K which was consistent with the previous experiments in air. The combination of these observations suggests that the oxidation rate is decreasing with temperature above 2500K. This phenomenon is well known; however, it is the first such observation for an ablating body with the additional influence of a full realistic hypervelocity shock layer. Improved magnification in the high-speed video recordings for these experiments also allowed spallation phenomena to be observed in much greater detail than in previous work. It was clearly seen that spallation can significantly alter the flowfield, and is a phenomenon which must be given greater consideration in future studies.