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A series of experiments were conducted in a F109 turbofan engine to investigate the unsteady wake profiles of an Inlet Guide Vane (IGV) at a typical spacing to the downstream fan at subsonic and transonic relative blade velocities. The sharp trailing-edge vanes were designed to produce a wake profile consistent with modern IGV. Time averaged baseline measurements were first performed with the IGV located upstream of the aerodynamic influence of the fan. Unsteady experiments were performed with an IGV-fan spacing of 0.43 fan chords. High-frequency on-vane pressure measurements showed strong peak-to-peak amplitudes at the blade passing frequency (BPF) of 4.7 psi at the transonic fan speeds. High-frequency total pressure measurements of the IGV wake were taken between the IGV and fan. Results showed that the total pressure loss coefficient of the time averaged IGV wake is reduced by 30% for the subsonic fan, and increased by a factor of 2 for the transonic fan compared to the baseline. Time resolved wake profiles for subsonic fan speeds show constructive and destructive interactions over each blade pass generated by the fan potential flow field. Time resolved wake profiles for the transonic fan speeds show that shock interactions with the IGV surface result in the wake shedding off of the vane at the BPF. Furthermore, the effectiveness of trailing edge blowing (TEB) flow control was investigated. TEB is the method of injecting air aft of the IGV to reduce the low pressure regions (deficits) in the viscous wakes shed by the vanes. Minimizing the IGV wakes reduces the forcing function on the downstream fan blades, thereby reducing high cycle fatigue. The TE span of the vane contains discrete holes at the axial centerline for TEB. Baseline results showed that TEB eliminates the IGV wake, while using only 0.03% of the total engine mass flow per IGV. TEB for the subsonic fan at the close spacing shows complete wake filling using the same mass flow as the baseline. TEB for the transonic fan shows a reduction of 68% in the total pressure loss coefficient, while requiring 2.5 times the mass flow as the baseline. / Ph. D.

One way to increase cycle efficiency of a gas turbine engine is to operate at higher turbine inlet temperature (TIT). In most engines, the turbine inlet temperatures have increased to be well above the metallurgical limit of engine components. Film cooling of gas turbine components (blades and vanes) is a widely used technique that allows higher turbine inlet temperatures by maintaining material temperatures within acceptable limits. In this cooling method, air is extracted from the compressor and forced through internal cooling passages within turbine blades and vanes before being ejected through discrete cooling holes on the surfaces of these airfoils. The air leaving these cooling holes forms a film of cool air on the component surface which protects the part from hot gas exiting the combustor. Design optimization of the airfoil film cooling system on an engine scale is a key as increasing the amount of coolant supplied yields a cooler airfoil that will last longer, but decreases engine core flow—diminishing overall cycle efficiency. Interestingly, when contemplating the physics of film cooling, optimization is also a key to developing an effective design. The film cooling process is shown to be a complex function of at least two important mechanisms: Increasing the amount of coolant injected reduces the driving temperature (adiabatic wall temperature) of convective heat transfer—reducing heat load to the airfoil, but coolant injection also disturbs boundary layer and augments convective heat transfer coefficient due to local increase in freestream turbulence. Accurate numerical modeling of airfoil film cooling performance is a challenge as it is complicated by several factors such as film cooling hole shape, coolant-to-freestream blowing ratio, coolant-to-freestream momentum ratio, surface curvature, approaching boundary layer state, Reynolds number, Mach number, combustor-generated high freestream turbulence, turbulence length scale, and secondary flows just to name a few. Until computational methods are able to accurately simulate these factors affecting film cooling performance, experimental studies are required to assist engineers in designing effective film cooling schemes. The unique contribution of this research work is to experimentally and numerically investigate the effects of coolant injection rate or blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane at high freestream turbulence (Tu = 16%) and engine representative exit flow conditions. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. The same facility was also used to conduct experimental and numerical study of the effects of freestream turbulence, and Reynolds number on smooth (without film cooling holes) turbine blade and vane heat transfer at engine representative exit flow conditions. The showerhead film cooled vane was instrumented with single-sided platinum thin film gauges to experimentally determine the Nusselt number and film cooling effectiveness distributions over the surface from a single transient-temperature run. Showerhead film cooling was found to augment Nusselt number and reduce adiabatic wall temperature downstream of injection. The adiabatic effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition region at all blowing ratio and exit Mach number conditions. The experimental study was also complimented with a 3-D CFD effort to calculate and explain adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of a turbine vane at high freestream turbulence (Tu = 16%) and engine design exit flow condition (Mex = 0.76). The research work presents a new three-simulations technique to calculate vane surface recovery temperature, adiabatic wall temperature, and surface Nusselt number to completely characterize film cooling performance in a high speed flow. The RANS based v2-f turbulence model was used in all numerical calculations. CFD calculations performed with experiment-matched boundary conditions showed an overall good trend agreement with experimental adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of the vane. / Ph. D.

The main contributor to the high-cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds is especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced blade vibrations was evaluated in a modern single-stage transonic compressor rig. A row of wake generator (WG) vanes with TEB capability was installed upstream of the fan blisk, the blades of which were instrumented with strain gages. Data was collected for varied TEB conditions over a range of rotor speed which included one fundamental and multiple harmonic resonance crossings. Sensitivity of resonant response amplitude to full-span TEB flowrate, as well as optimal TEB flowrates, are documented for multiple modes. Resonant response sensitivity was generally characterized by a robust region of substantial attenuation, such that less-than-optimal TEB flowrates could prove to be an appropriate design tradeoff. The fundamental crossing amplitude of the first torsion mode was reduced by as much as 85% with full-span TEB at 1.1% of the total rig inlet flow. Similar reductions were achieved for the various harmonic crossings, including as much as 94% reduction of the second leading edge bending mode resonant response using 0.74% of the rig flow for full-span TEB. At least 32% reduction was achieved for all modal crossings over the broad flow range of 0.5 to 0.9% of the rig flow. Thus the results demonstrate the modal- and flowrate-robustness of full-span TEB for reducing forced response in a modern, closely-spaced transonic compressor. Reduced spanwise TEB coverage was generally found to provide less peak reduction. Widely varying sensitivities of the vibration modes to the spanwise TEB distribution were also noted. While the second chordwise mode experienced roughly the same maximum response reduction of 80% for all of the spanwise TEB configurations, some other modes were amplified from the baseline case under part-span TEB conditions. Part-span TEB was thus found to be less modally-robust than full-span TEB. / Ph. D.

A comprehensive study of the flowfield through a two-dimensional cascade of the high pressure turbine blades of a jet engine is presented. The main interest is the measurement and prediction of the mass-averaged total pressure losses. Other experiments, such as flow visualization, are aimed at the validation of the code that was used to obtain the numerical results and also to further knowledge about the details of the loss generation. The experimental studies were carried out on a cascade of eleven blades in a blow-down tunnel. Total pressure measurements were taken upstream of the cascade and also by traversing on downstream planes. The static pressures needed for the mass averaging and the probe bow shock correction were obtained by pressure taps on the cascade tunnel side wall. The static pressure was also measured on the surface of some instrumented blades. Shadowgraph pictures were taken for study of the trailing edge shock structure and for the turbulent transition location. A single-plate interferometer technique was used for density field measurements. The major goal of the numerical studies was the prediction of the mass-averaged total pressure losses, but all other measured quantities were also generated from the computed flowfield. A critical issue was the generation of a proper grid. For the studied type of flow, a non-periodic C-type grid turned out to be the most advantageous. For use in the moderately compressible attached turbulent boundary layer, a Clauser-type eddy viscosity model was developed and tested. In the trailing edge and wake region, the Baldwin-Lomax model was used. Good agreement of calculations and measurements was obtained for the blade surface and cascade tunnel side wall static pressures, the trailing edge shock structure, and the density field. The agreement between the measured and calculated total pressure drop profiles was not quite as good; however, that quantity is known to be difficult to predict accurately. The mass-averaged total pressure loss coefficient, calculated from the total pressure drop profiles, was again in good agreement with the measurements. The difference between the measured and computed total pressure drop profiles suggested that the Baldwin-Lomax model underpredicted the eddy viscosity in the trailing edge region. / Ph. D.

The contribution of present research work is to experimentally investigate the effects of blowing ratio and mainstream Mach number/Reynolds number (from 0.6/8.5X10⁵ to 1.0/1.4X10⁶) on the performance of the fan-shaped hole injected turbine blade and vane. The study was operated with high freestream turbulence intensity (12% at the inlet) and large turbulence length scales (0.26 for blade, 0.28 for vane, normalized by the cascade pitch of 58.4mm and 83.3mm respectively). Both convective heat transfer coefficient, in terms of Nusselt number, and adiabatic effectiveness are provided in the results. Present research work also numerically investigates the shock/film cooling interaction. A detailed analysis on the physics of the shock/film cooling interaction in the blade cascade is provided. The results of present research suggests the following major conclusions. Compared to the showerhead only vane, the addition of fan-shaped hole injection on the turbine Nozzle Guide Vane (NGV) increases the Net Heat Flux Reduction (NHFR) 2.6 times while consuming 1.6 times more coolant. For the blade, combined with the surface curvature effect, the increase of Mach number/Reynolds number results in an improved film cooling effectiveness on the blade suction side, but a compromised cooling performance on the blade pressure side. A quick drop of cooling effectiveness occurs at the shock impingement on the blade suction side near the trailing edge. The CFD results indicate that this adiabatic effectiveness drop was caused by the strong secondary flow after shock impingement, which lifts coolant away from the SS surface, and increases the mixing. This secondary flow is related to the spanwise non-uniform of the shock impingement. / Ph. D.

In this thesis an approximate method is developed which predicts the aerodynamic force on a cone-cylinder body in axially symmetric transonic now. The method places more emphasis on the physics of the now than on the mathematical rigors of solving the typical reduced non-linear transonic equation of motion. Under the assumption that the now is that of a steady, irrotational, inviscid, compressible gas, the body pressures are determined and the associated force defined. Recognizing that the transonic pressures are influenced by the character of the subsonic compressible pressures, which are obtained in this analysis through Gothert’s Rule, it is then mandatory that the incompressible case be defined with the best possible accuracy. Comparisons with experiments indicate that the classical method (axially distributed sources and sinks) does not provide this required accuracy. Thus the surface distributed vortex ring theory is used in the present analysis to obtain the incompressible body pressures. Gothert’s Rule, which represents a linear solution for the subsonic case, is known to be applicable up to a limit value of tree stream Mach number. An investigation is carried out herein to determine both the correct form of the rule and its limits of applicability. As a result of this investigation, it is concluded that the upper limit is the lower free stream critical Mach number. Also, at this Mach number, a solution is immediately available tor the lower limit of the transonic range of Mach number. In solving the transonic problem the law or stationarity of local Mach number is of fundamental importance. For an assumed isentropic flow over the body, and for sonic conditions being present at some point on the surface, the body pressures can be described in the ratio p_L/p*. Here p_L is the local surface pressure and p* is the sonic (body) pressure. Through the stationarity law, this ratio is recognized as an invariant for transonic speeds so long as the flow field remains essentially irrotational. Thus any change in local pressure is only a function of the free stream Mach number for any given body position. By this approach, the pressure distribution is defined for a range of Mach number from below to above the sonic stream value. The method is then capable of prediction for almost all of the transonic range of Mach number. It is only when the head shock baa significant curvature, causing the now adjacent to the body to be rotational, that the method fails. Though the procedure developed here is not capable of spanning the entire transonic range, it does provide a wider range of applicability than other known theories. Finally, for this problem, a correlation of transonic pressure drag data is formulated. This correlation is founded on physical interpretation and is not limited to the usual transonic similarity restrictions. In fact, to the author's knowledge, this is the first known such correlation tor axially symmetric flow covering the range of body sizes and Mach numbers considered in this investigation. In so far as is practicable the results obtained in this thesis have been compared to available experimental results. In particular, the drag data from this analysis compare closely with experimental transonic values. Experiment bears out the conclusion that the upper limit for linear theory is the lower critical tree stream Mach number. And, the pressures determined by the vortex ring theory agrees well with the low-speed experimental results obtained by the author. / Ph. D.

An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a 1st stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/1,400,000; 0.85/1,150,000; and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/1,150,000 three blowing ratio conditions were tested: BR = 1.0, 1.5, and 2.0. At a Mach/Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR = 1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by leading edge diameter of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5x) over the showerhead only NGV and agreement with published showerhead-shaped hole data. NHFR was shown to increase substantially (average 2.6x increase) with the addition of shaped holes, with only a small increase (average 1.6x increase) in required coolant mass flow. Heat transfer and effectiveness augmentation with increasing blowing ratio was shown on the pressure side, however the suction side was shown to be less sensitive to changing blowing ratio. Boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number. / Master of Science

This study utilizes Multidisciplinary Design Optimization (MDO) techniques to explore the effectiveness of the truss-braced (TBW) and strut-braced (SBW) wing configurations in enhancing the performance of medium range, transonic transport aircraft. The truss and strut-braced wing concepts synergize structures and aerodynamics to create a planform with decreased weight and drag. Past studies at Virginia Tech have found that these configurations can achieve significant performance benefits when compared to a cantilever aircraft with a long range, Boeing 777-200ER-like mission. The objective of this study is to explore these benefits when applied to a medium range Boeing 737-800NG-like aircraft with a cruise Mach number of 0.78, a 3,115 nautical mile range, and 162 passengers. Results demonstrate the significant performance benefits of the SBW and TBW configurations. Both configurations exhibit reduced weight and fuel consumption. Configurations are also optimized for 1990's or advanced technology aerodynamics. For the 1990's technology minimum TOGW cases, the SBW and TBW configurations achieve reductions in the TOGW of as much as 6% with 20% less fuel weight than the comparable cantilever configurations. The 1990's technology minimum fuel cases offer fuel weight reductions of about 13% compared to the 1990's technology minimum TOGW configurations and 11% when compared to the 1990's minimum fuel optimized cantilever configurations. The advanced aerodynamics technology minimum TOGW configurations feature an additional 4% weight savings over the comparable 1990's technology results while the advanced technology minimum fuel cases show fuel savings of 12% over the 1990's minimum fuel results. This translates to a 15% reduction in TOGW for the advanced technology minimum TOGW cases and a 47% reduction in fuel consumption for the advanced technology minimum fuel cases when compared to the simulated Boeing 737-800NG. It is found that the TBW configurations do not offer significant performance benefits over the comparable SBW designs. / Master of Science

The effects of a shock wave passing through a blade passage on surface heat transfer to turbine blades were measured experimentally. The experiments were performed in a transonic linear cascade which matched engine Reynolds number, Mach number, and shock strength. Unsteady heat flux measurements were made with Heat Flux Microsensors on both the pressure and suction surfaces of a single blade passage. Unsteady static pressure measurements were made using Kulite pressure transducers on the blade surface and end walls of the cascade. The experiments were conducted in a stationary linear cascade of blades with heated transonic air flow using a shock tube to introduce shock waves into the cascade. A time-resolved model based on conduction in the gas was found to accurately predict heat transfer due to shock heating measured during experimental tests without flow. The model under-predicted the experimental results with flow, however, by a factor of three. The heat transfer increase resulting from shock passing in heated flow averaged over 200 its (typical blade passing period) was found to be a maximum of 60% on the pressure surface near the leading edge. Based on experimental results at different flow temperatures, it was determined that shock heating has the primary effect on heat transfer, while heat transfer increase due to boundary layer disturbance is small. / Master of Science