Matthew McGilvray

Analysis of air radiation measurements obtained in the EAST and X2 shocktube facilities

This paper presents measurements of equilibrium radiation obtained in the NASA Ames Research Centers EAST facility and the University of Queenslands X2 facility. These experiments were aimed at measuring the level of radiation encountered during conditions relevant to Orion lunar return into Earths atmosphere. The facilities have targeted the same nominal test conditions of 10 km/s and 26.6 Pa (0.2 Torr). In addition, variations on the nominal shock speed have also been the focus of recent testing in the EAST facility. A comprehensive comparison between the EAST data and NEQAIR is presented in this paper with preliminary X2 comparisons where appropriate. Since the two facilities have different dimensions, and the tests have different shock speeds, NEQAIR simulations are used as a point of reference for the EAST and X2 comparison. Results obtained by independently reducing the data from both facilities are compared. The present analysis endeavors to provide a better understanding of the uncertainty in the measurements, as well as provide an initial comparison between EAST and X2. Furthermore, the present analysis explores various radiative mechanisms to determine if they are due to physical processes relevant to flight, or are just facility dependent phenomena. These phenomena include effects such as the magnitude of the background continuum.

Scramjet Experiments in an Expansion Tunnel: Evaluated Using a Quasi-Steady Analysis Technique

As scramjet engine development moves toward higher and higher speeds, it will be important to continue to ground test complete nose-to-tail configurations at true flight conditions. Above Mach 10, the freestream total pressure requirements can only be met by an expansion-tube facility. To establish the practicality of ground-testing complete scramjet configurations at high Mach numbers, an expansion-tunnel facility is used to test a generic two-dimensional scramjet at a Mach 10 replication condition. As the flow produced in an expansion-tube facility is transient in nature, a method for analyzing point measurements within the scramjet is developed. A transient numerical simulation of the engine configuration without fuel injection is used to verify the validity of the technique, as well as to provide a reference data set for comparison to experimental data. These experiments show that stable supersonic combustion is established during the test time. Using the developed transient analysis, accurate quasi-steady pressure data at true flight conditions can be inferred from scramjet test measurements in an expansion-tube facility.

Expansion tunnel radiation experiments to support Hayabusa re-entry observations

The Hayabusa sample return capsule is scheduled for re-entry near Woomera, Australia in June 2010 and expansion tube experiments are being performed to support the planned re-entry observation campaign. Initial experiments using a 1/10th scale model of the Hayabusa forebody have been performed in the X2 expansion tunnel facility at The University of Queensland to simulate aerothermal elements of the anticipated re-entry. Experiments have been performed at an effective flight speed of around 9.8 km/s using steel models, and steel models coated with a layer of epoxy to simulate pyrolysis gases associated with heat shield ablation. Spectral emissions from the stagnation region of the capsule have been acquired using a spectrograph system. Two dimensional maps of the luminous emissions from the shock heated flow have also been acquired using a high speed camera. Deduction of flow conditions generated in the X2 expansion tunnel is achieved using quasione-dimensional simulations coupled to an axisymmetric simulation of the flow through the expansion tunnel nozzle. The effects of the ablative epoxy material are observed in the data from both the spectrograph system and the high speed camera. Both systems register strong emissions in the ablative layer, and the strength of the spectral peaks associated with CN emissions are shown to be enhanced by the presence of the epoxy. Further measurement and analysis is required to confidently define the flow conditions produced by the expansion tunnel, and to quantify results from the spectrograph and high speed camera measurements.

Comparison of Mach 10 Scramjet Measurements from Different Impulse Facilities

A comparison is made between scramjet pressure data taken in two different impulse facilities, a reflected shock tunnel and an expansion tunnel, to provide verification of flow similarity. A two-dimensional hydrogen-fueled scramjet was tested in both facilities at a Mach 10, 66 kPa dynamic pressure, flight-replication condition. The entire flowfield was reproduced, including the inlet, combustor, and thrust nozzle. The expansion-tunnel model was a 40%-scaled model of that investigated in the shock tunnel, as the generated core flow was smaller. Applying density-length scaling, the density of the freestream was increased by a factor of 2.5 in the expansion tube to maintain similarity between the experimental setups. Results show agreement within the experimental uncertainty between the two data sets at similar fuel equivalence ratios, indicating the repeatability of scramjet pressure measurements, independent of the experimental facility used. It is further demonstrated that density-length scaling is appropriate for the flows generated in the investigated scramjet engines. Verification is also provided of the analysis technique used to process experimental scramjet data sets that involve transient inflow conditions. Copyright

Effects of Upstream Injection on Scramjet Performance Using an Entropy-Based Method

Upstream injection has the potential for increasing the overall performance of a scramjet by reducing the overall length required to complete combustion and, thereby, reducing overall vehicle frictional drug. However, high flow losses will be incurred by injecting a fuel on the inlet where the Mach number is high. A simple overall approach based on entropy generation was used to investigate the effect or flow losses oil overall performance. This showed that the injection flow losses can he quite high for upstream injection when compared with conventional combustor injection. An inviseid analysis of a two-dimensional scramjet at a Mach 10 flight condition found that a significant loss In overall performance of the engine occurs if the bulk of injection is undertaken on the intake. Estimates of viscous effects in a scramjet combustor showed that depending on the reduction of the mixing length achieved in the combustor, the upstream injection flow losses could he substantially recovered, making it an advantageous design option. The analysis technique developed, like similar simple analysis methodologies, provides a useful understanding of the physical processes that influence thrust production and can also be used to quickly evaluate and optimize potential configurations before starting detailed flow-field simulations.

T6: The Oxford University Stalker Tunnel

The University of Oxford has embarked on developing the UKs fastest wind tunnel, T6, in collaboration with the University of Queensland (Australia) using technology pioneered by the late Professor Ray Stalker. The T6 facility couples the ex-Australian National University T3 free piston driver to the barrels, nozzles and test section of the Oxford gun tunnel. The facility can operate in three different modes; as a shock tube, a reected shock tunnel or as an expansion tunnel. The T6 facility will be unique to Europe allowing for hypervelocity ground testing not possible in any current EU facilities, whilst having the exibility to conduct tests across a large range of speeds and binary scaling products (pL) of interest.

Boundary Layer Transition in an Expansion Tube at a Low Enthalpy Operating Condition

A full flight similarity condition for scramjet testing in an expansion tube was seen ex-perimentally to have a large drop in pitot pressure midway through the test time. This be-haviour has been studied experimentally, numerically and analytically and was determined to be caused by boundary layer transition on the wall of the expansion tube. Compari- son of numerical calculations and experimental measurements of heat transfer on the wall have shown a good agreement of both heat transfer level and transition location, although the numerical simulation does not exhibit the same extent of boundary layer growth after transition. Three possible ways to mitigate the effects of this problem have been suggested while still matching all flow parameters during the test time. These are (1) the use of a gas other than air in the acceleration tube, (2) the use of a reflected shock at the end of the shock tube and (3) the use of the steady nozzle on the end of the acceleration tube. So far, the analysis has shown that use of the steady expansion nozzle will relieve the problems of core flow interruption due to boundary layer transition and allow testing of scramjets in an expansion tube at true flight conditions. This solution also gives the benefit of a longer test time and larger core flow diameter possible with the straight tube configuration. AIAA 2007-1328 Copyright

Radiation measurements in nonreflected shock tunnels

Atmospheric entry at superorbital speeds is associated with large amounts of radiation, often in nonequilibrium modes. This is a significant design issue for applications such as lunar return vehicles, aerocapture and for all missions to the gas giants. Analyzing such flows is complicated by a lack of both physical measurements from shock heated gases, and by our current understanding of the fundamental nonequilibrium processes involved. Shock tubes provide a useful means of obtaining realistic data from radiating gases under the same conditions as encountered in flight, and represent a powerful tool for the validation of advanced radiation models. This paper reports on work in progress to adapt expansion tubes to operate in the nonreflected shock tube mode to extend the range over which such data can be usefully obtained. The results of preliminary tests in the X3 facility are shown, and the possibilities for extending the measurements to a larger flight envelope are discussed.

Reduction in Flow Parameter Resulting From Volcanic Ash Deposition in Engine Representative Cooling Passages

Internal cooling passages of turbine blades have long been subject to blockage through the deposition of sand and dust during fleet service life. The ingestion of high volumes of volcanic ash therefore poses a real risk to engine operability with the additional difficulty that the cooling system is frequently impossible to inspect to assess the level of deposition. This paper reports results from experiments carried out at typical HP turbine blade metal temperatures (1163K to 1293K) and coolant delivery temperatures (800K to 900K) in engine scale models of a turbine cooling passage with film-cooling offtakes. Volcanic ash samples from the Eyjafjallajokull eruption were used for the majority of the experiments conducted. Ash particle size distributions typical of those reaching internal cooling systems were generated through the sieving of the ash, with subsequent measurement of the distribution and characterization of the particle shape through SEM analysis. A further ash sample from the Chaiten eruption allowed the effect of changing ash chemical composition to be investigated. The experimental rig allows the metered delivery of volcanic ash through the coolant system at the start of a test. The geometry, rate of injection, metal and coolant temperatures can be independently varied, allowing the sensitivity of passage blockage to each parameter to be determined. The key metric indicating blockage is the flow parameter which can be determined over a range of pressure ratios (1.01 – 1.06) before and after each experiment. The data are reported at hot and cold conditions. Visual inspection of the test pieces is useful in determining where deposition has occurred. Results from the experiments have determined the threshold metal temperature at which blockage occurs for the ash samples available, and characterise the reduction of flow parameter with changing particle size distribution, blade metal temperature, ash sample composition, film-cooling hole configuration and pressure ratio across the holes. There is qualitative evidence that hole geometry can be manipulated to decrease the likelihood of blockage. A discrete phase CFD model implemented in Fluent has allowed the trajectory of the ash particles within the coolant passages to be modelled, and these results are used to help explain the behaviour observed.

Computational and Experimental Study of Heat Transfer in Rotating Ribbed Radial Turbine Cooling Passages

This study investigates the effect of rotation on the Nusselt number distribution within ribbed radial turbine cooling passages representative of systems used in current jet engines. The results are unusual in that the cooling passage length to diameter ratio is engine representative and full distributions of local Nusselt number have been measured using the transient liquid crystal method. The results are compared to RANS CFD simulations and the level of agreement discussed in detail. A triple-pass serpentine passage is investigated, which includes 45° filleted rib-turbulators and 180° curved bends. The first two passes have an aspect ratio of 1:4 which are radially inward and outward respectively, with the final pass being radially outward with an aspect ratio of 1:2. The Reynolds, Rotation and Buoyancy numbers are all representative of a passage within a HP turbine blade of a gas turbine engine at 97000/108000, 0.081/0.088 and 0.052/0.035 respectively for the 1:4/1:2 aspect ratio passages. CFD simulations are found to give good predictions under stationary conditions however significant differences are observed when rotation is introduced. The Nusselt number distributions depend strongly on both rotation and upstream flow conditions created by the specific geometry.Copyright