Development of a Fast-Response Calorimeter Gauge for Hypersonic Ground Testing

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.

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.

A new method of measuring heat transfer rate with high spatial resolution and fast response is developed to be used in wind tunnels with long flow duration. For the development of the new method a multi-layered thin film heat transfer gauge is proposed. The principle of the sensor is based on measuring the temperature gradient across a thin, heat resistant layer of SiO with two, thin film resistance thermometers on its upper and lower surfaces. These thin films and the heat resistant layer are deposited by vacuum evaporation technique. The design consideration, the accuracy of the sensor, frequency response, and the calibration technique are discussed. The sensor is applied to the measurement of the heat transfer rate in turbulent boundary layer on a flat plate at Mach number 4, Tw/T0= 0.64, and Re=1.4×107, and shows high sensitivity and fast response. It also shows excellent durability.

The impulse response method is widely used for heat transfer analysis in turbomachinery applications. Traditionally, the 1D method assumes a: linear time invariant, isotropic, semi-infinite block with planar surfaces and does not accurately model the true geometric behaviour. This paper evaluates the error introduced by the planar assumption and outlines the required modifications for accurate freeform surface analysis. Adapted cylindrical basis functions are defined for the impulse response method and used to evaluate the impact of the 1D planar assumption. The analytic solutions for both convex and concave surfaces are presented. A penta-diagonal algorithm, for a modified numerical Crank-Nicolson scheme, is also evaluated for fast stable implementation of curved geometry simulation. The scheme shows comparable performance to the impulse response, whilst removing the requirement for linear time invariance. The new methods are demonstrated in the case of aerothermal analysis for heat transfer in a turbine nozzle guide vane. A 3D ANSYS simulation is used as a benchmark and further highlights the importance of the curvature effects. The methods are extended using curvature mapping for infrared camera data, enabling direct 3D thermal simulation or the fast calculation of freeform curvature. This paper defines the methodology to analyse heat transfer measurements on non-planar geometry.

Hypervelocity Wind Tunnel No. 9, located at the White Oak, MD site of the Arnold Engineering Development Complex (AEDC), has long been recognized as a unique world class ground-test facility. The facility was developed in the early 1970s to provide critical low-altitude, high Mach number data in support of the Navy's reentry development programs. Since its inception, Tunnel 9 has maintained a leading role in hypersonic ground testing by continually expanding its operational capabilities to match the needs of current and projected programs, maintaining data quality, and understanding customer requirements. Tunnel 9 started with a unique design built around a state-of-the-art supply heater that provided a clean, high-pressure, high-temperature nitrogen supply. Initial operation of Tunnel 9 realized a Mach 10 and 14 aerodynamic simulation capability. Additional Mach 7 and 8 high Reynolds number capabilities were subsequently developed. Each upgrade to Tunnel 9 during the past 40 years of operation has been in response to various sponsors or hypersonic basic research requirements. These capability enhancements have helped maintain Tunnel 9's position as a core DoD hypersonic test and evaluation (T&E) ground-test facility, which has been identified as a leading facility in all major hypersonic facility studies. Recent improvements and modernization over the past 10 years have focused on test article measurements and have significantly changed the types, quantity, and quality of test data that are readily acquired in a Tunnel 9 test entry. Recent advancements include major system changes such as a completely new control room and high-speed data system to entirely new measurement capabilities such as global heattransfer measurements using Temperature Sensitive paint technology. These along with other incremental improvements have allowed Tunnel 9 to provide new insights into the physics associated with complex hypersonic flows.

In this paper, experiments to detect turbulent spots in the transitional boundary layers, formed on a flat plate in a free-piston shock tunnel flow, are reported. Experiments indicate that thin-film heat-transfer gauges are suitable for identifying turbulent-spot activity and can be used to identify parameters such as the convection rate of spots and the intermittency of turbulence.

Hypersonic airliner would be exposed to temperatures that are beyond the limits of classical aircraft materials. In order to handle this problem the latest developments of new materials and composite structures suitable for high temperature application need to be taken into account. The focus of the European Research program ATLLAS is on advanced light-weight, high-temperature material development strongly linked to a high-speed passenger aircraft design. ATLLAS stands for Aerodynamic and Thermal Load Interactions with Lightweight Advanced Materials for High Speed Flight. The 4 years program ATLLAS-II is a logical continuation project built upon the experience and technology development gained within ATLLAS-I. The corresponding work related to combustor structures and material integration deals with the opportunity to investigate at academic level, both in basic and relevant environment, different solutions possibly usable to ensure the long range cruise of a high speed airliner. Different materials (UTHC, CMC, metallic) and different cooling techniques (radiation, convective, transpiration) are studied. Available numerical or semi-empirical tools are used to prepare the test, to design the different architectures. A pin fin experiment allows to better know the pressure drop and the heat transfer for different channel patterns with thermal cristal techniques. The ERBURIG-K long duration test facility allow to characterize different ceramic matrix composite uncooled samples and will allow to realize, at small scale, a long duration (several hours) investigation of cooled ceramic structure in PTAH-SOCAR technology. A multifonctionnal metallic transpiration cooled HSS panel using Hollow Spheres Stacking as core material was designed and preliminary tested in cold conditions with GN2 and in hot conditions with infra lamps under 1 MW/m² heat flux. CMC and UHTC materials are used to design, manufacture and test generic fin injectors usable in high speed combustors. Industrial hypersonic METHYLE test facility is used to test in relevant Mach 6 combustor environment HSS panel as well as advanced fin injectors. Hot and cold permeability of composites is documented with GN2 and GH2. Numerical models are used in accordance with the experiments, some examples are given in the present paper.

A current R&D program at AEDC plans to explore advanced technologies to develop a hypersonic test facility, followed by a multiyear prototype facility development effort. The AEDC Strategic Planning Office (AEDC/XPX) has an effort underway to define the future test conditions that will define the performance requirements for an aeropropulsion hypersonic test facility. The Boeing Phantom Works is under contract to provide information on projected hypersonic vehicles and develop test requirements for them. The first phase defined advanced hypersonic tactical missile, hypersonic cruise, and space access vehicles that may be developed within 15 to 20 years with airbreathing engines. Size, shape, weight, aerodynamic and propulsion performance, and trajectories were defined for four baseline vehicle concepts. The first phase of the study provided an analysis of the ground test facility requirements for run time, model size, Mach number, enthalpy, and pressure for performance, operability, and durability testing. The second phase has provided an analysis of the ground test facility flow quality requirements by conducting sample calculations and discussing contamination effects, chemical kinetics, vibrational relaxation, and condensation issues. Our analyses in this study indicated that the presence of combustion products in the freestream flow would not seriously undermine the validity of test results and conclusions for testing at Mach 8 and above for typical test conditions. Test facilities that use combustion heating vitiation to simulate high Mach conditions can produce useful results but require careful posttest analysis. Our analyses in this study also indicated that the presence of NOX species did not cause significant test or posttest analysis difficulties.

: The much ignored problem of turbulence in hypersonic ground test facilities is discussed with particular emphasis on arc-heated facilities. Sources of free-stream turbulence and their effects on model aerodynamics are identified. An analytical treatment of stagnation zone heating accounting for free-stream vorticity amplification and viscous dissipation is presented. Results are discussed in relation to three experiments: one involving time resolved measurements in an arc-heated wind tunnel, another involves stagnation point heat transfer measurements in the turbulent boundary layer of a hypersonic shock tunnel, and the third deals with similar measurements in a turbulent subsonic nitrogen plasma jet. In both of the latter experiments, stagnation point heating rates exceeding two to three times the calculated laminar expectation are reported. These results are corroborated by the present analysis.

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.

An evaluation of analog and numerical techniques for unsteady heat transfer measurement with thin-film gauges in transient facilities

Vehicles entering planetary atmospheres at high speed require an ablative heat shield to withstand the high thermal energy flux to the body. The interaction between the ablative products and the flowfield is not well characterized. In this study, numerical simulations were conducted to investigate the influence of carbon ablation on shock-layer radiation. Data collected from experiments performed in the X-2 expansion tunnel at the University of Queensland were used to compare to the simulations. The model was a short half-cylinder made of isomolded graphite and was tested in Earth entry flow. The graphite model was heated within a temperature range of 1770–2410 K over the course of the experimental campaign. The radiation emitted from the CN violet bands was measured by ultraviolet spectrometry in a spectral range from 353 to 391 nm. The simulations used the Park as well as Zhluktov and Abe finite-rate surface kinetic models for determining the chemical state of the ablating boundary layer. The Park model consistently overpredicted the radiative heat fluxes, but better comparison was achieved when a modified nitridation rate was used. The Zhluktov and Abe predictions displayed good agreement with the measured radiative heat fluxes at 1770 K but remained well below the measured values at the higher wall temperatures. The lack of a nitridation mechanism in the Zhluktov and Abe model is suggested as the reason for these underpredictions.

Numerical modelling of thin film gauge for short duration thermal measurement

Augmentation of hypersonic propulsion test facilities using MHD

Results from initial operation of a continuous-flow, arc-heated hypersonic propulsion test facility