<p>  </p> <p>As aviation fuel costs and consumption have continued to rise over recent decades, gas turbine engine manufacturers have sought methods to reduce fuel burn. Manufacturers plan to achieve this by reducing the specific fuel consumption of the machine by increasing the bypass ratio through a reduction of the diameter of the engine core. This presents an opportunity for implementing a centrifugal compressor as the final stage of the high-pressure compressor. The vaned diffuser in a centrifugal compressor stage maintains an integral role in determining the extents of the operating range as well as conditioning the flow for the downstream combustor. Thus, it is critical to understand the aerodynamics and performance of the vaned diffuser across the entire compressor operating range.</p> <p>This investigation focused on seven compressor operating points at the stage’s design corrected speed, which ranged from choked flow to the minimum mass flow rate before rotating stall. Steady-state and unsteady performance data were acquired to study the aerodynamics at each operating point as well as the steady-state performance of the vaned diffuser. Laser Doppler velocimetry, high-frequency pressure transducers, and additive manufacturing techniques were all implemented to acquire data in the research compressor.</p> <p>Unsteady velocity measurements were acquired in the vaneless space and were used to quantify the change in diffuser inlet incidence as the stage mass flow rate changes. The impeller exit jet and wake were compared at each operating point to understand the effect of these flow structures on the spanwise incidence profile. Steady-state performance metrics for the vaned diffuser were compared with the change in incidence to assess the effect of incidence on performance. Maximum static pressure recovery and minimum total pressure loss occurred at the maximum incidence operating point. </p> <p>The chordwise static pressure distribution is critical for health monitoring of the polymer, additive manufactured diffuser vanes. Steady-state and unsteady pressure measurements were acquired along the diffuser vane surface to assess the change in the aerodynamic lift force across the compressor operating range as well as the static pressure differential across the vane leading edge. The maximum unsteady lift on the diffuser vanes was up to 34% greater than the steady-state lift force. Unsteady static pressure differentials across the diffuser vane leading edge were similar to the steady-state values, but they were marginally greater across the entire examined operating range. These unsteady pressure measurements were acquired with high-frequency response pressure transducers installed along the diffuser vane surfaces. These transducers were also used to study the rotating stall and surge behavior of the investigated centrifugal compressor stage. This centrifugal compressor stage exhibits a spike-type rotating stall pattern at the onset of stage instability, which rapidly evolves into full flow reversal with several surge cycles. During these surge cycles, the diffuser vane leading edges are subject to a 20 psid static pressure differential. </p> <p>A computational model was used to predict the compressor flow at three different operating points. This model utilized the BSL-EARSM turbulence model, and it included surface roughness and an experimentally measured shroud thermal profile. The model accurately predicted the diffuser inlet flow angles near the shroud, but it predicted more radial flow near midspan. The diffuser vane leading edge static pressure differential was predicted within 1 psid at higher aerodynamic loading conditions. The differences between the computationally predicted and experimentally measured flow are attributed to difficulties associated with modelling the rate of mixing within the flow.</p>

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<p>Highly loaded airfoils in low-pressure turbines (LPTs) suffer from laminar flow separation from the suction side of the airfoils aft of the throat of the passages. This separation harms the performance of the engine by reducing the power extraction from the turning air and ultimately reduces the overall turbine efficiency. Flow control techniques have been investigated to eliminate flow separation in aerodynamic surfaces to abate the losses associated with it. This Master of Science Thesis investigates the design, implementation and testing of pulsated injection actuation in a low-Reynolds flow over a wall-mounted hump.</p> <p>Furthermore, this Thesis expands on the existing expertise in the infrared (IR) thermography measurement technique at the Purdue Experimental Turbine Aerothermal Lab. This is done through an investigation of the factors affecting the IR measurement technique and the development of an optical instrument (borescope) to implement in an annular cascade wind tunnel. IR thermography is used on the wall-mounted hump blowdown tests to detect the separation point in the boundary layer using two techniques: by an investigation of the surface temperature distribution and an investigation of the heat transfer behavior at the surface. Finally, the borescope is commissioned through the first testing campaign of the LPT airfoils, and are processed to thermally investigate the passage.</p> <p>This thesis succeeds in expanding the IR capabilities within PETAL, and at demonstrating pulsated injection as an effective method to eliminate flow separation. Furthermore, IR successfully detects flow separation on the wall-mounted hump through the two methods presented, as well as detecting the boundary layer reattachment caused by the flow control technique. The limitations of the thermal methodology, as well as those of the optical probe are addressed, and the uncertainties in the measurements are quantified. Finally, steps to continue the studies are suggested at the end of each methodology chapter, including the potential redesign of the IR borescope to improve the quality of measurements. </p>

Turbine engine efficiency modeling depends on many parameters related to fluid dynamics and heat transfer. Many of these parameters change dynamically once the engine enters service and begins to experience surface degradation. This thesis presents a validation of the design and operation of an accelerated testing facility for the study of foreign deposit layers typical to the operation of land-based gas turbines. It also reports on the use of this facility in an effort to characterize the change in thermal resistance on the surface of turbine blades as deposits accumulate. The facility was designed to produce turbine deposits in a 4-hour test that would simulate 10,000 hours of turbine operation. This is accomplished by matching the net foreign particulate throughput of an actual gas turbine. Flow Mach number, temperature and particulate impingement angle are also matched. Validation tests were conducted to model the ingestion of foreign particulate typically found in the urban environment. The majority of this particulate is ceramic in nature and smaller than 10µm in size, but varies in size up to 80µm. Deposits were formed for flow Mach number and temperature of 0.3 and 1150°C respectively, using air plasma sprayed (APS) thermal barrier coat (TBC) material coupons donated from industry. These conditions are typical of a modern, first stage nozzle. Investigations over a range of impingement angles yielded samples with deposit thicknesses from 50 to 200µm in 4-hour, accelerated-service simulations. Above a threshold temperature, deposit thickness was dependent primarily upon particle concentration. Test validation was achieved using direct comparison with deposits from service hardware. Deposit characteristics affecting blade heat transfer via convection and conduction were assessed. Surface topography analysis indicated that the surface structure of the generated deposits were similar to those found on actual turbine blades. Scanning electron microscope (SEM) and x-ray spectroscopy analyses indicated that the deposit microstructures and chemical compositions were comparable to turbine blade deposit samples obtained from industry. A roadmap for the development of a theoretical model of thermal resistance using the SEM scan is presented. Thermal resistance experiments conducted with deposit samples indicate that a general decrease in thermal resistance occurs as the samples are exposed to operating conditions in the accelerated testing facility. This is likely due to sintering effects within the TBC dominating any thermal resistance increase arising from deposition. Recommendations for future research into the interaction between TBC sintering and deposit evolution are presented.

Multiple high-fidelity, time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single stage transonic compressor. Various configurations of a transonic axial compressor stage, including three stator/rotor axial spacings of close, mid, and far in conjunction with three stator loadings of decreased, nominal, and increased were simulated in order to understand the flow physics of transonic blade-row interactions. Low-speed compressors typically have reduced stator/rotor axial spacing in order to decrease engine weight, and also because there is an increase in efficiency with reduced axial spacing. The presence of a rotor bow shock in high-speed compressors causes additional losses as the shock interacts with the upstream stator trailing edge. This research analyzes the strength of shock-induced vortices due to these unsteady blade-row interactions. The time-accurate URANS code, TURBO, was used to generate periodic, quarter annulus simulations of the Blade Row Interaction compressor rig. Both time-averaged and time-accurate results compare well with experimentally-observed trends. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Normal and large shock-induced vortices formed on the stator trailing edge immediately after the shock passing, but the large vortices were strengthened at the trailing edge due to a low-velocity region on the suction surface. This low velocity region was generated upstream of mid-chord on the suction surface from a shock-induced thickening of the boundary layer or separation bubble, due to the rotor bow shock reflecting off the stator trailing edge and propagating upstream. The circulation of the shock-induced vortices increased with shock strength (decreased axial spacing) and stator loading. Most design tools do not directly account for unsteady effects such as blade-row interactions, so a model is developed to help designers account for shock-induced vortex strength with varying shock strength and stator loading. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics early in the design process.

This paper examines two examples of testing activities of sub-components of a 300kN methalox first-stage rocket engine developed by Pangea Aerospace, namely, theinjectors and the cryogenic pumps. One of the difficulties of these test campaignsis to evaluate the performances of these sub-components without using cryogenicpropellants and in conditions differing from flight conditions.Two different methods were used to replace the cryogenic propellants at extreme pressures and temperatures: the pumps were tested using water to replace both propellants, and the injectors were tested using water and nitrogen to account for the different propellant phases. The correlation, similarity methods, and bench sizing were detailed for both test campaigns, but only the injector test campaign was completed and its results were analyzed.The method to emulate cryogenic fluid behavior in a pump leads to the use of a bench controlling not only pressure and mass flow but also temperature to account for the cavitation effects. Furthermore, the bench must have the capability to pressurize the water under atmospheric pressure, leading to the use of a vacuum pump. The injector test campaign produces results that closely align with theoretical predictions. This enables the selection of the optimal single-element injector design through a quantitative assessment of the discharge coefficient and a qualitative evaluation of the spray angle and atomization. Although these studies are based on strong models, they must be compared to hotfire data and later flight data to be assessed fully, all the more so that the pump test campaign was not performed. / I denna rapport behandlas två exempel på provningsaktiviteter för delkomponenter av en methalox-driven 300 kN-klass raketmotor, utvecklad av Pangea Aerospace för användning på förstasteg. Delkomponenterna består av injektorerna samt dekryogena pumparna. En av svårigheterna med provkampanjerna är att utvärdera prestandan hos dessa delkomponenter utan att använda kryogena drivmedel och underförhållanden som skiljer sig från flygförhållandena. Två olika metoder användes för att imitera de kryogena drivmedlen vid extrema tryck och temperaturer. Först testades pumparna med vatten som en ersättning för de båda drivmedlen, och sedan testades injektorerna med både vatten och kväve för att ta hänsyn till de olika aggregations tillstånden. Korrelationen, likhetsmetoderna, samt provbänkstorleken utredes i detalj för bägge testkampanjer, men endast injektortestkampanjen slutfördes och dess resultat analyserades. Metoden för att efterlikna kryogena vätskors beteende i en pump leder till användning av en provbänk som inte bara reglerar tryck och massflöde utan även temperatur för att ta hänsyn till kavitations effekterna. Dessutom måste bänken ha kapacitet att trycksätta vattnet under atmosfärstryck, vilket leder till användning av en vakuumpump. Injektortestkampanjen ger resultat som ligger nära de teoretiska förutsägelserna. Detta gör det möjligt att välja den optimala injektordesignen genom en kvantitativ bedömning av utloppskoefficienten och en kvalitativ bedömning av vinkeln och finfördelningen av injektorstrålen. Även om dessa studier baseras på starka modeller måste de jämföras med data från brännprover och senare flygdata för att kunna bedömas fullt ut, särskilt som pumptestkampanjen inte genomfördes.

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The work presented in this thesis aims to provide various tools to be used during design process to make maximum use of the increasing availability of accurate engine blade measurement data for high fidelity fluid mechanic simulations at a reasonable computational expense. A new method for uncertainty propagation for geometric error has been proposed for fluid mechanics codes using adjoint error correction. Inexpensive Monte Carlo (IMC) method targets small uncertainties and provides complete probability distribution for the objective function at a significantly reduced computational cost. A brief literature survey of the existing methods is followed by the formulation of IMC. An example algebraic model is used to demonstrate the IMC method. The IMC method is extended to fluid mechanic applications using Principal Component Analysis (PCA) for reduced order modelling. Implementation details for the IMC method are discussed using an example airfoil code. Finally, the IMC method has been implemented and validated for an industrial fluid mechanic code HYDRA. A consistent methodology has been developed for the automatic generation of the linear and adjoint codes by selective use of automatic differentiation (AD) technique. The method has the advantage of keeping the linear and the adjoint codes in-sync with the changes in the underlying nonlinear fluid mechanic solver. The use of various consistency checks have been demonstrated to ease the development and maintenance process of the linear and the adjoint codes. The use of AD has been extended for the calculation of the complete Hessian using forward-on-forward approach. The complete mathematical formulation for Hessian calculation using the linear and the adjoint solutions has been outlined for fluid mechanic solvers. An efficient implementation for the Hessian calculation is demonstrated using the airfoil code. A new application of the Independent Component Analysis (ICA) is proposed for manufacturing uncertainty source identification. The mathematical formulation is outlined followed by an example application of ICA for artificially generated uncertainty for the NACA0012 airfoil.