Skin Friction Sensor for High-Speed, High-Enthalpy Scramjet Flow Applications

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 Relation between flow structure and local skin friction on the wall surface of the clustered linear aerospike nozzle models was investigated by conducting cold flow experiments using a skin friction sensor which has been developed in this study. The skin friction sensor is capable of measuring the local skin friction value on the wall. Measured skin friction value was compared to some existing prediction model and results found that the local skin friction value becomes drastically different from the one predicted by existing models. In cases of over-expansion condition skin friction was drastically increased by oblique shock wave. In cases of under-expansion condition skin friction was increased which was caused by interference of jet boundary of cell-nozzle flows jetted from adjoined cellnozzle.

This study presents an application of the naphthalene sublimation technique to examine the manner in which solid fuels utilized in air-breathing engines may be expected to regress under various flow conditions. A counterflow burner rig was used to exhibit the increase in the sublimation rate of naphthalene when subjected to greater levels of shear stress and surface temperatures. Additionally, skin friction sensor data was taken using Ahmic Aerospace’s supersonic wind tunnel, with a naphthalene coated flat plate instrumented with three skin friction sensors installed in the wall of the wind tunnel. It was found that wall shear decreased an average of 11% across the three sensors over the course of four Mach 2.4 wind tunnel runs. This indicates that naphthalene sublimation can be used as an effective tool to study changes in surface roughness and skin friction at supersonic Mach numbers. Future work with Ahmic’s skin friction sensors and thermocouples embedded within cast naphthalene will be conducted in Virginia Tech’s Hypersonic Wind Tunnel at the Advanced Power and Propulsion Laboratory. This will allow for further definition of the skin friction experienced by naphthalene, which can be used to characterize the skin friction experienced by solid fuels in similar flow scenarios. The naphthalene sublimation technique provides analogous data to study the rate and manner of regression of solid fuels which can also be utilized to refine computer models of surface regression.

A skin friction sensor is a three-dimensional MEMS sensor specially designed for measuring the skin friction of hypersonic vehicle models. The accuracy of skin friction measurement under hypersonic laminar flow conditions is closely related to the fabrication and micro-assembly accuracy of MEMS skin friction sensors. In order to achieve accurate skin friction measurement, high-precision linear laser scanning ranging, multi-axis precision drive, and 3D reconstruction algorithms are investigated; a MEMS skin friction sensor micro-assembly height error measurement system is developed; and the MEMS skin friction sensor micro-assembly height error control method is carried out. The results show that the micro-assembly height error measurement of MEMS skin friction sensors achieves an accuracy of up to 2 μm. The height errors of the MEMS skin friction sensor were controlled within −8 μm to +10 μm after error control. The angular errors were controlled within the range of 0.05–0.25°, significantly improving micro-assembly accuracy in the height direction of the MEMS skin friction sensor. The results of hypersonic wind tunnel tests indicate that the deviation in the accuracy of the MEMS skin friction sensors after applying height error control is about 5%, and the deviation from the theoretical value is 8.51%, which indicates that height error control lays the foundation for improving the accuracy of skin friction measurement under hypersonic conditions.

The skin friction of a hypersonic vehicle surface can account for up to 50% of the total resistance, directly affecting the vehicle’s effective range and load. A wind tunnel experiment is an important and effective method to optimize the aerodynamic shape of aircraft, and Micro-Electromechanical System (MEMS) skin friction sensors are considered the promising sensors in hypersonic wind tunnel experiments, owing to their miniature size, high sensitivity, and stability. However, the sensitive structure including structural appearance, a gap with the package shell, and flatness of the sensor will change the measured flow field and cause the accurate measurement of friction resistance. Aiming at the influence of sensor-sensitive structure on wall-flow characteristics and friction measurement accuracy, the two-dimensional and three-dimensional numerical models of the sensor in the hypersonic flow field based on Computational Fluid Dynamics (CFD) are presented respectively in this work. The model of the sensor is verified by using the Blathius solution of two-dimensional laminar flow on a flat plate. The results show that the sensor model is in good agreement with the Blathius solution, and the error is less than 0.4%. Then, the influence rules of the sensitive structure of the sensor on friction measurement accuracy under turbulent flow and laminar flow conditions are systematically analyzed using 3D numerical models of the sensor, respectively. Finally, the sensor-sensitive unit structure’s design criterion is obtained to improve skin friction’s measurement accuracy.

Uncertainty quantification (UQ) is the study of the effects of uncertainty on the values of analytical results and the predictions of scientific models. Sources of uncertainty include imprecise knowledge of the exact values of parameters, lack of confidence in the physical models, use of imperfectly calibrated models, and irreducible uncertainties due to physical characteristics. The Air Force Research Laboratory has undertaken the challenge of understanding, developing and analyzing the techniques of UQ as they apply to Laser Beam Control. This paper proposes a simple methodology and simple results with our first attempt of applying UQ as a new analysis tool. The software toolkit which was chosen was an analytical group of algorithms from a Sandia National Laboratory (SNL) package called DAKOTA (Design Analysis Kit for Optimization and Terascale Applications). The specific application of interest to the Air Force Research Laboratory (AFRL) is the analytical prediction of the performance of a Laser Beam Control systems under various scenarios, conditions, and missions. The application of rigorous UQ techniques to the models used to predict beam control performance could greatly improve our confidence in these predictions and also improve the acceptance of advanced Laser Beam Control systems within the science and engineering communities1,2. The proposed work would follow a multi-step approach, analyzing the more easily quantified sources of uncertainty, and then including increasingly complicated physical phenomena as the work progresses. Will present the initial results, and the first steps in the incorporation of UQ into our Laser Beam Control Modeling and Simulation environments.

An experimental investigation was carried out to measure skin friction and heat flux in a Mach 6 freejet test of a scramjet. inlet component. The investigation was performed in the Air Force Research Laboratory’s Mach 6 windtunnel; The model simulated a movable cowl mounted on a sidewall compression plate. The skin friction gauge‘used’a floating element supported by a miniature cantilever beam which is instrumented with strain gauges to measure the beam deflection. Coaxial thermocouples were used to measure the wall temperature from which the heat flux was.deduced. Several measurements.were taken on the sidewall compression plate without the cowl lip, effectively,,a flat plate. This information was then used to .determine the boundary layer transition point and the influence on the boundary layer profile. The data is useful both for drag assessment of scramjet inlets and for CFD ‘turbulence model development. Comparison between the measured skin friction .and two turbulent boundary layer skin friction correlations, the Van Driest II and Spalding and Chi, showed that the Spalding and Chi more accurately matched the measured conditions. A Reynolds analogy factor of 1.29 is required to match skin friction data to heat flux measurements.

Mass spectrometry is a versatile research tool that has proved to be extremely useful for exploring the fundamental nature of the earth's atmosphere and ionosphere and in helping to solve operational problems facing the Air Force and the Department of Defense. In the past 40 years, our research group at the Air Force Research Laboratory has flown quadrupole mass spectrometers of many designs on nearly 100 sounding rockets, nine satellites, three Space Shuttles and many missions of high-altitude research aircraft and balloons. We have also used our instruments in ground-based investigations of rocket and jet engine exhaust, combustion chemistry and microwave breakdown chemistry. This paper is a review of the instrumentation and techniques needed for space research, a summary of the results from many of the experiments, and an introduction to the broad field of atmospheric and space mass spectrometry in general.

This study investigates the optimal placement of sensors mounted over the flap of a twodimensional (2D) airfoil equipped with Coanda jet circulation control for the purpose of flow control. The analysis is based on flow field data obtained from 2D unsteady Reynolds averaged Navier-Stokes (uRANS) simulations of a high-lift airfoil with different jet blowing intensities and actuation frequencies, characterizing different flow separation states. Linear stochastic estimation of the proper orthogonal decomposition (POD) is used to estimate the time-coefficients. The optimal sensor locations for any number of sensors configuration are then determined by minimizing the root mean squared error between the determined and the estimated time-coefficients. The results are compared for two types of sensor inputs: pressure sensors and skin-friction sensors. The results show that it is possible to determine the state of flow more accurately using fewer pressure sensors placed at optimized locations than from using the same number of skin friction sensors.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

The Propulsion Systems Lab (PSL) at the NASA Glenn Research Center contains two large engine te st cells, PSL -3 and PSL -4, capable of simulating flight conditions up to 70,000 ft with a maximum forward air speed of Mach 3.0. The facility has experienced high levels of inlet temperature distortion at high altitude flight conditions. This report desc ribes how the inlet temperature profiles were improved for future jet engine testing. Temperature distortion levels were improved from 7 percent to 1.8 percent by modifying the cold air supply piping system. Flow angle and turbulence intensity were measu red in the plenum, upstream of the engine inlet, to investigate possible sources of distortion. As a result, plenum flow quality and turbulence intensity for PSL4 are documented in this report. I. Introduction he Propulsion Systems Lab (PSL) at the NASA Gle nn Research Center contains two large engine test cells, PSL -3 and PSL -4, capable of simulating flight conditions up to 70,000 ft with a maximum forward air speed of Mach 3.0. Recently, PSL -4 has been used for high altitude testing of small engines. High altitude test conditions are typically low speed and require cooled inlet air at temperatures down to -40 degrees F. The facility has experienced high levels of inlet temperature distortion at these simulated flight conditions. The temperature distortio ns are much higher than would be experienced during actual flight. This report summarizes an investigation into high levels of temperature distortion in the PSL -4 inlet plenum. The temperature distortion can cause undesirable effects on jet engine test ar ticles, such as reduced surge margin. Reduced surge margin can cause unstable test conditions, or can adversely affect experiments to determine surge margin. This report describes how the inlet temperature profiles were improved for future jet engine tes ting.

A theoretical and experimental study of an ionizing laminar boundary layer formed by a very high enthalpy flow (in excess of 12 eV per atom or 7000 cal/gm) with allowance for the presence of helium driver gas is described. The theoretical investigation has shown that the use of variable transport properties and their respective derivatives is very important in the solution of equilibrium boundary layer equations of high enthalpy flow. The effect of low level helium contamination on the surface heat transfer rate is minimal. The variation of ionization is much smaller in a chemically frozen boundary layer solution than in an equilibrium boundary layer calculation and consequently, the variation of the transport properties in the case of the former was not essential in the integration. The experiments have been conducted in a free piston shock tunnel, and a detailed study of its nozzle operation, including the effects of low levels of helium driver gas contamination has been made. Neither the extreme solutions of an equilibrium nor of a frozen boundary layer will adequately predict surface heat transfer rate in very high enthalpy flows.

Excitation of carbon monoxide molecules has been carried out in a cold cell and in a low-pressure plasma jet using an ArF narrow-band excimer laser. The different excitation models are discussed and the relevance of atomic carbon absorption into the laser cavities is pointed out. Excitation spectra of Cameron bands have been obtained in a room-temperature cell and compared with calculated spectra. A value of the constant σ related to the interaction strength between a 3 Π( v = 2) state and its neighbouring singlet states is derived: 0≤σ≤0.05. The fluorescence spectrum following broad band excitation of CO has been observed both in UV and visible. Similar experiments carried out in a high enthalpy flow have allowed to point out the presence of a 3 Π metastable carbon monoxide. A method for relative measurements of this species concentration is proposed.

A RCJET thrusters have been developed throughout the past few decades. They provide relatively high thrust at moderate exhaust velocities, and their plasma flows are characterized by high specific enthalpy and high flow velocity. These characteristics are, however, combined with the presence of steep radial property gradients as in a hot, energy-rich core with a relatively cold gas layer at its edge. When the limits to the nozzle exit velocity are imposed bywall and electrode materials, another possibility to transfer more power into an arc heated plasma flow is to reheat the relatively cold plume edges by means of another heating mechanism, as done in an afterburner. TIHTUS has been developed over the past four years and is a novel two-stage plasma thruster where reheating of the arc heated plume is realized bymeans of inductive heating, as sketched in Fig. 1 [1–3]. It is therefore considered the predecessor of a future propulsion system for the transport of heavy payload on interplanetary trajectories [1]. The principal question within the development of this thruster is whether it is possible to specifically heat the outer edges of an arcjet plume so that higher exhaust velocity can be attained. A dependency of the power staging between the two thruster stages and of the gas mass flow rate staging is expected. For the presented research, the thruster was investigated with a cavity calorimeter which provides data of plasma power so that a total efficiency can be derived. Moreover, a method for deducing the two-stage system’s efficiency from operational data is presented and its data are compared with the measured calorimetric data.

As a method for analysis of supersonic flow, visualization of the flow field structure using LIF of tracers seeded in a carrier gas has been employed. In this study, the capability of imaging of supersonic flow using O2-LIPF (laser-induced predissociative fluorescence) is examined. Since oxygen is one of the main constituents of air, O2-LIPF may enable visualization of air flow in a high-speed wind tunnel without tracers. However, because the visualization of the flow field below room temperature using O2-LIPF is very difficult, it has been applied only to high enthalpy and combustion flow, not to high-speed flow at low temperature. In this study, the supersonic oxygen flow is visualized by O2-LIPF using an ArF excimer laser and a high-sensitivity CCD camera. If the CCD camera is set at a long exposure time (1200 sec) and the scattered laser beam is suppressed sufficiently, the supersonic oxygen flow below room temperature can be imaged successfully using O2-LIPF.