An approximate method of calculating the minimum total weight of the two-insulation--two-coolant thermal protection system is developed. The equations derived in the development of the approximate method enable insight into the parameters that control the system weight. Two cases are considered: the case where the outer coolant location is unrestricted within the insulating wall and the case where the outer coolant location is restricted within the insulating wall. The effects on system weight of material properties and the outer coolant location within the insulating wall are discussed. A comparison of weights predicted by the approximate method and numerical solutions is shown. / Master of Science

The identification of linear structural dynamic systems has been dealt with extensively in past studies. Identification methods for nonlinear structures have also been introduced in previous articles, including procedures based on the method of multiple scales, iterative and noniterative direct methods, and state space mappings. Here, a procedure is introduced for the identification of nonlinear structural dynamic systems which is readily applicable to simple as well as more complex multiple degree of freedom systems. The procedure is based on multiple step integration methods for the solution of differential equations. The multiple step integration procedure and the iterative direct method are applied to a number of nonlinear single degree of freedom examples, and are applied to a simple two degrees of freedom example as well. RMS based noise is added to a simulated measured response in order to monitor the effects of measurement errors on the procedures. The input data is filtered before final processing in the identification algorithms. The multistep algorithm is compared to the iterative direct method on the basis of criteria such as accuracy, ease of use, and numerical efficiency. / Master of Science

A study was conducted to determine the effect of strut position on the aerodynamic and acoustic performance of a supersonic inlet. The investigated inlet was a prototype 1/14 scale model of a mixed compression, axisymmetric supersonic inlet designed for the high speed civil transport aircraft. A 10.4 cm (4.1 in.) turbofan engine simulator was used in conjunction with the inlet to provide the typical noise signature of a high bypass turbofan engine. Two inlet configurations were investigated in this study. The first configuration was the standard inlet design where the struts are located immediately upstream of the fan. The new configuration has the struts located 3.3 chord length upstream of the fan. The purpose for relocating the strut position was to reduce the flow distortion and radiated noise level. The experiment was conducted at various fan operating conditions in order to simulate aircraft approach. The inlet was tested statically without simulating the inflight speed effects. Steady state measurements were made in order to evaluate the aerodynamic performance of the inlet. The aerodynamic results show that the movement of the struts to a new location allowed the strut wake to diffuse significantly before reaching the fan. This reduced the circumferential distortion parameter by a factor of 2.4, without affecting the pressure recovery of the inlet at a fan Abstract speed of 30,000 rpm (40 PNC). Acoustic measurements were taken in the far field in the 0°-110°sector from the inlet axis. The new configuration of the inlet showed an improved acoustic performance over the standard design. Relocating the struts upstream reduced the blade passing tone by an average of 8 dB (0°-110°) sector, and the overall sound pressure level was lowered by an average of 2.6 dB at a fan speed of 30,000 rpm (40 PNC). / Master of Science

The use of aerodynamic forces to assist in certain orbital transfers can greatly reduce the fuel consumption as compared with corresponding all-propulsive transfers. Therefore, in seeking minimum fuel trajectories, aeroassisted transfers need to be investigated. A review of the current literature indicates that such problems have been solved almost exclusively via optimal control theory formulations that result in continuously varying control laws. The use of a piecewise-constant strategy allows the controls to vary to a degree necessary to affect changes in the desired state dependent parameters while simplifying the optimization process. In the process of searching for a tool to produce numerical results, the current research investigates three candidate methods of solving the parameter optimization problem of minimum fuel aeroassisted orbital transfer with piecewise-constant controls. A method based on implicitly integrating the state trajectory is chosen over methods which analytically and explicitly integrate the state trajectory. The implicit method offers improved performance over the explicit method while presenting a more correct solution than the analytic method. The analytic method is shown to suffer from approximations that lead to undesirable solutions. Analytic expressions for the characteristic velocities of Hohmann and idealized aeroassisted transfers are presented and compared. For a large number of transfers from high Earth orbit to low Earth orbit, the aeroassisted mode requires less fuel. Numerical results are presented for minimum fuel transfer from geosynchronous Earth orbit to low earth orbit for a variety of control strategies. The piecewise-constant strategies are seen to provide solutions which are comparable to those found via optimal control theory. / Master of Science

This investigation concerns the design, build, and testing of a new class of skin friction sensor capable of performing favorably in high-speed, high-enthalpy flow conditions, such as that found in atmospheric re-entry vehicles, scramjets, jet engines, material testing, and industrial processes. Fully understanding and optimizing these complex flows requires an understanding of aerodynamic properties at high enthalpies, which, in turn, requires numerical and analytical modeling as well as reliable diagnostic instrumentation. Skin friction is a key quantity in assessing the overall flight and engine performance, and also plays an important role in identifying and correcting problem areas. The sensor design is founded on a direct-measuring, cantilever arrangement. The design incorporates two fundamental types of materials in regards to thermal conductivity and voltage resistivity properties. The non-conducting material distinction greatly deters the effect of heat soak and prevents EMI transmission throughout the sensor. Four custom fabricated metal-foil strain gauges are arranged in a Wheatstone bridge configuration to increase sensitivity and to provide further compensation for sensitivity effects. The sensor is actively cooled via a copper water channel to minimize the temperature gradient across the electronic systems. The design offers a unique immunity to many of the interfering influences found in complex, high-speed, high-enthalpy flows that would otherwise overshadow the desired wall shear measurement. The need to develop an encompassing design methodology was recognized and became a principal focus of this research effort. The sensor design was developed through a refined, multi-disciplinary approach. Concepts were matured through an extensive and iterative program of evolving key performance parameters. Extensive use of finite element analysis (FEA) was critical to the design and analysis of the sensor. A software package was developed to utilize the powerful advantage of FEA methods and optimization techniques over the traditional trial and error methods. Each sensor endured a thorough series of calibrations designed to systematically evaluate individual aspects of its functionality in static, dynamic, pressure, and thermal responses. Bench-test facilities at Virginia Tech (VT) and Air Force Research Laboratory (AFRL) further characterized the design vibrational effects and electromagnetic interference countermeasure effectiveness. Through iterations of past designs, sources of error have been identified, controlled, and minimized. The total uncertainty of the skin friction sensor measurement capability was determined to be ±8.7% at 95% confidence and remained fairly independent of each test facility. A rigorous, multi-step approach was developed to systematically test the skin friction sensor in various facilities, where flow enthalpy and run duration were progressively increased. Initial validation testing was conducted at the VT Hypersonic Tunnel. Testing at AFRL was first performed in the RC-19 facility under high-temperature, mixing flow conditions. Final testing was conducted under simulated scramjet flight conditions in the AFRL RC-18 facility. Performance of the skin friction sensors was thoroughly analyzed across all three facilities. The flow stagnation enthalpies upward of 1053 kJ/kg (453 Btu/lbm) were tested. A nominal Mach 2.0 to 3.0 flow speed range was studied and stagnation pressure ranged from 172 to 995 kPa (25 to 144 psia). Wall shear was measured between 94 and 750 Pa (1.96 and 15.7 psf). Multiple entries were conducted at each condition with good repeatability at ±5% variation. The sensor was also able to clearly indicate the transient flow conditions of a full scramjet combustion operability cycle to include shock train movement and backflow along the isolator wall. The measured experimental wall shear data demonstrated good agreement with simple, flat-plate analytical estimations and historic data (where available). Numerical CFD predictions of the scramjet flow path gave favorable results for steady cold and hot flow conditions, but had to be refined to handle the various fueling injection schemes with burning in the downstream combustor and surface roughness models. In comparing CFD wall shear predictions to the experimental measurements, in a few cases, the sensor measurement was adversely affected by shock and complex flow interaction. This made comparisons difficult for these cases. The sensor maintained full functionality under sustained high-enthalpy conditions. No degradation in performance was noted over the course of the tests. This dissertation research and development program has proven successful in advancing the development of a skin friction sensor for applications in high-speed, high-enthalpy flows. The sensor was systematically tested in relevant, high-fidelity laboratory environments to demonstrate its technology readiness and to successfully achieve a technology readiness level (TRL) 6 milestone. The instrumentation technology is currently being transitioned from laboratory development to the end users in the hypersonic test community. / Ph. D.

This thesis aims at presenting the Discrete Vortex Method (DVM) as a tool to determine the flow field and associated wind loads over structures. Two structures are considered: the first is a surface-mounted prism and is used to simulate wind loads over low-rise structures. The second is a bridge section with attached flaps that can be oriented to vary the moment coefficient. Advantages and disadvantages of using DVM for these applications are discussed. For the surface-mounted prism, the results show that the developed code correctly predicts the flow separation around the corners. As for the surface pressures, it is concluded that parallel processing, which could be easily implemented for DVM, should be used to correctly predict surface pressures and their variations. This is due to the required slow time advancement of the computations. The results on attaching flaps to bridge sections yield required orientations to minimize moments under different angles of attack. / Master of Science

Morphing aircraft combine both radical and subtle wing shape changes to improve vehicle performance relative to a rigid airframe. An aircraft wind tunnel model with considerable wing-shape freedom can serve as a tool in learning to model, control, and fully exploit the potential of such vehicles. This work describes the design, development, and initial analysis of a wind tunnel model that combines large and small wing shape variations for fundamental research in modeling and control of morphing air vehicles. The vehicle is designed for five primary purposes: quasi-steady aerodynamic modeling of an aircraft with large planform changes, optimization studies in achieving efficient flight configurations, transient aerodynamic modeling of high-rate planform changes, evaluating planform maneuvering as an control effector, and gimbaled flight control simulation of a morphing aircraft. The knowledge gained from the wind tunnel evaluations will be used to develop general stabilization and optimal control strategies that can be applied to other vehicles with large scale planform changes and morphing flight models. After a brief background on the development of the Morphing Aircraft Program, and previous research ventures, the first phase vehicle development is described. The vehicle function, subsystems, and control are all presented in addition to the results of first phase wind tunnel testing. Deficiencies in the phase one design motivated the phase two development which has led to the current vehicle model: MORPHEUS. The evolution towards the MORPHEUS configuration is presented in detail along with an elementary strength analysis. The new embedded control implementation to permit a rate controllable planform is included. A preliminary aerodynamic analysis is presented to contrast MORPHEUS against the phase one design and an industry morphing concept. In particular, it is shown how the redesigned model has enhanced performance characteristics and the additional degrees of freedom enable greater flexibility in optimizing a configuration, especially with respect to trim characteristics. An expansion of traditional analysis techniques is applied to derive a new optimal twist algorithm for the MORPHEUS model at each planform configuration. The analysis concludes with a hybrid continuous modeling method that combines first-order computational aerodynamic modeling with classic stability expressions and DATCOM enhancements. The elementary aerodynamic coefficients are computed over the range of possible planform configurations and combined with the optimal twist results for preliminary trim analysis. This work precedes phase two wind tunnel testing and transient modeling. Future work involves expansion into the five purposes detailed for the MORPHEUS model. / Master of Science

The possibility of misalignments at interfaces would be increased due to individual parts' assembly and external factors during its operation. In actual engine representative conditions, the upstream misalignments have effects on turbines performance through the nozzle guide vane passages. The current experimental aerodynamic investigation over the nozzle guide vane passage was concentrated on the backward-facing step of upstream misalignments. The tests were performed using two types of vane endwall platforms in a 2D linear cascade: flat endwall and axisymmetric converging endwall. The test conditions were a Mach number of 0.85, Re_ex 1.5*10^6 based on exit condition and axial chord, and a high freestream turbulence intensity (16%), at the Virginia tech transonic cascade wind tunnel. The experimental results from the surface flow visualization and the five-hole probe measurements at the vane-passage exit were compared with the two cases with and without the backward-facing step for both types of endwall platforms. As a main source of secondary flow, a horseshoe vortex at stagnation region of the leading edge of the vane directly influences other secondary flows. The intensity of the vortex is associated with boundary layer thickness of inlet flow. In this regard, the upstream backward-facing step as a misalignment induces the separation and attachment of the inlet flow sequentially, and these cause the boundary layer of the inlet flow to reform and become thinner locally. The upstream-step positively affects loss reduction in aerodynamics due to the thinner inlet boundary layer, which attenuates a horseshoe vortex ahead of the vane cascade despite the development of the additional vortices. And converging endwall results in an increase of the effect of the upstream misalignment in aerodynamics, since the inlet boundary layer becomes thinner near the vane's leading edge due to local flow acceleration caused by steep contraction of the converging endwall. These results show good correlation with many previous studies presented herein. / Master of Science / In response to climate change and limited resources, fossil fuel prices are expected to rise and energy policies are expected to change. Under these circumstances, there is a growing demand in the industry to provide an affordable option for improving the efficiency of technology. Energy efficiency is one of most cost effective ways to improve the competitiveness of all businesses and reduce energy costs for consumers. Regarding the current study topic in particular, the gas turbine is an internal combustion engine that extracts energy, which is resultant from the liquid fuel flow, and is then converted into mechanical energy to drive a compressor or other devices. Gas turbines are used in many applications such as, to power aircraft, electrical generators, pumps, and gas compressors in industrial fields. Because the gas turbine has a probability of unaligned connections of components due to assembly characteristics of its huge size, performance is affected. To consider issue, an experimental study was conducted related to the energy efficiency for an actual engine’s representative conditions; the current study focuses on the upstream backward facing step of the unaligned connections and highlights the practical effects of the unaligned connection and converging geometry. These backward facing unaligned connections are shown to have positive effects for reducing aerodynamic losses by weakening a main source of the loss, even despite the development of the additional losses. And, the application of converging geometry to the gas turbine also results in loss reduction due to local flow acceleration. These results show good correlation with the many previous studies presented herein.

In this thesis, we describe a method for controlling the cycle-averaged velocity direction of a fixed-wing aircraft in an unpowered, helical descent. While the aircraft propulsion system is disabled, either intentionally or due to a failure, the aerodynamic control surfaces (aileron, elevator, and rudder) are assumed to be functional. Our approach involves two steps: (i) establishing a stable, steady, helical motion for which the control surfaces are not fully deflected and (ii) modulating the aircraft control surfaces about their nominal positions to ``slant'' the helical flight path in a desired direction relative to the atmosphere, whether to attain a desired impact location, to counter a steady wind, or both. The effectiveness of the control law was evaluated in numerical simulations of a general transport model (GTM). / Master of Science / When an unmanned aircraft is near an authorized airspace (a region of space where the aircraft is not authorized to fly) and experiences a failure such as loss of communication with the control tower,or failure of the GPS or propulsion system, then the aircraft is generally put into an aerodynamic flight termination. In this flight termination method, the aircraft propulsion system is switched off and the control surfaces (aileron, elevator and rudder) are fixed to induce a spin in the aircraft causing it to descend in a helical fashion. However, in the presence of external gusts the aircraft might drift into the unauthorized airspace; or once the aircraft is put into spin, one may want to be able to change the impact location to a safer place. To the best of our knowledge, there exist no control strategies to alter the impact location of the aircraft once it is put into spin and while is continues to spin. In this thesis we describe a method to do so. The aircraft impact location is altered by controlling the cycle-averaged velocity direction of a fixed-wing aircraft in an unpowered, helical descent. While the aircraft propulsion system is disabled, either intentionally or due to a failure, the aerodynamic control surfaces (aileron, elevator, and rudder) are assumed to be functional. Our approach involves two steps: (i) establishing a stable, steady, helical motion for which the control surfaces are not fully deflected and (ii) modulating the aircraft control surfaces about their nominal positions to “slant” the helical flight path in a desired direction relative to the atmosphere, whether to attain a desired impact location, to counter a steady wind, or both. The effectiveness of the control law was evaluated in numerical simulations of a general transport model (GTM).

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