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This paper describes various techniques employed in a novel approach to instruction and assessment of an undergraduate sequence in thermo-propulsion at the United States Air Force (USAF) Academy. Integrated motivational contexts aid development of foundations in thermodynamics, compressible gas dynamics, and propulsion while reinforcing engineering problem solving skills. Students are first oriented to the context of new material. Subsequent lessons fortify the context, giving students the opportunity to collaborate on team design projects and interact with industry and government guest speakers. Real-world, practical examples and homework further motivate and help students grasp key concepts. Tests are administered in both oral and written formats with open-ended, scenario-based questions to assess student understanding of fundamentals. Grading procedures focus on analytical methods as opposed to numerical results. Specific performance criteria validate the achievement of course educational outcomes. Student course critique scores and written comments further support the assertion that a contextual framework is highly effective in teaching fundamental thermo-propulsion concepts.

This paper describes the development and assessment of a sophomore-level, aero-thermodynamics class structured to meet the needs of both the Department of Aeronautics and Department of Astronautics at the United States Air Force Academy. The course was developed following ABET EC2000 guidelines. Because of the large core class requirement placed on students at the USAF Academy, this single course was developed as an alternative to students taking traditional separate thermodynamics and gas dynamics courses. Benefits and tradeoffs of this approach are presented. The general philosophy in developing the course was to provide solid foundations in thermodynamics and compressible gas dynamics while motivating and inspiring students to their chosen engineering profession. To that end, the course is loaded with practical applications and hands-on laboratories. Engineering rigor was maintained by inclusion of an unsteady, three-dimensional control volume formulation of the governing equations, emphasizing assumptions and their implications, and enforcing engineering analysis methods. Quantitative assessment of specific performance criteria demonstrates achievement of educational outcomes. Student course critique scores provided additional quantitative data. Finally, an initial assessment of course impact on two different undergraduate propulsion classes demonstrates the intended result — improved understanding of fundamentals allowing for expanded coverage in other areas. In short, the propulsion tracks in both departments appear to be improved.Copyright © 2005 by ASME

This paper describes a systematic and graphical approach for leading undergraduates or other inexperienced designers through the preliminary design process of rotating turbomachinery. It is based upon the general principles of classic thermal system design and optimization. The paper explains the steps required to arrive at a combination of design variables that yield a “design envelope” or workable domain of aerodynamic and mechanical solution space. Once the “design envelopes” for both the compressor and the turbine have been identified, the rotor angular velocities are matched. The students are introduced to the design concept of moving from the workable to the optimal using two Mathcad® 11 worksheets described in the paper. The process begins with obtaining the performance requirements and operating conditions form cycle analysis which leads to the selection of the aerodynamic design point. A system of equations based upon aerodynamic and mechanical principles is then optimized for both the compressor and the turbine. Two animated surface plots allow the designer to keep track of three independent design variables as they influence the value of an objective function that represents how close each design alternative is to achieving the target values of several important stage properties. Constraint maps reveal the range of acceptable rotor angular velocities for both the compressor and the turbine which enables a match to be made. The design approach is explained using a descriptive example of a multistage compressor and turbine operating on a high pressure spool. This paper provides a clear roadmap for systematically arriving at the correct combination of multiple design inputs that yield the appropriate multiple outputs required in the preliminary design of rotating turbomachinery.Copyright © 2004 by ASME

A streamline curvature (SLC) throughflow numerical model was assessed and modified to better approximate the flow fields of highly transonic fans typical of military fighter applications. Specifically, improvements in total pressure loss modeling were implemented to ensure accurate and reliable off-design performance prediction. The assessment was made relative to the modeling of key transonic flow field phenomena, and provided the basis for improvements, central to which was the incorporation of a physics-based shock loss model. The new model accounts for shock geometry changes, with shock loss estimated as a function of inlet relative Mach number, blade section loading (flow turning), solidity, leading edge radius, and suction surface profile. Other improvements included incorporation of loading effects on the tip secondary loss model, use of radial blockage factors to model tip leakage effects, and an improved estimate of the blade section incidence at which minimum loss occurs. Data from a single-stage, isolated rotor and a two-stage, advanced-design (low aspect ratio, high solidity) fan provided the basis for experimental comparisons. The two-stage fan was the primary vehicle used to verify the present work. Results from a three-dimensional, steady, Reynolds-averaged Navier-Stokes model of the first rotor of the two-stage fan were also used to compare with predicted performance from the improved SLC representation. In general, the effects of important flow phenomena relative to off-design performance of the fan were adequately captured. These effects included shock loss, secondary flow, and spanwise mixing. Most notably, the importance of properly accounting for shock geometry and loss changes with operating conditions was clearly demonstrated. The majority of the increased total pressure loss with loading across the important first-stage tip region was shown to be the result of increased shock loss, even at part-speed.

A streamline curvature method with improvements to key loss models is applied to a two-stage, low aspect ratio, transonic fan with design tip relative Mach number of approximately 1.65. Central to the improvements is the incorporation of a physics-based shock model. The attempt here is to capture the effects of key flow phenomena relative to the off-design performance of the fan. A quantitative analysis regarding solution sensitivities to model parameters that influence the key phenomena over a wide range of operating conditions is presented. Predictions are compared to performance determined from overall and interstage measurements, as well as from a three-dimensional, steady, Reynolds-averaged Navier-Stokes method applied across the first rotor. Overall and spanwise comparisons demonstrate that the improved model gives reasonable performance trending and generally accurate results. The method can be used to provide boundary conditions to higher-order solvers, or implemented within novel approaches using the streamline curvature method to explore complex engine-inlet integration issues, such as time-variant distortion.Copyright © 2002 by ASME

Two single-stage, transonic compressor designs were tested under various undistorted operating conditions to characterize the process leading up to aerodynamic stall. The rig case was instrumented with eight high-response static pressure transducers equally spaced around the annulus for stall development detection. High-response measurements were low-pass filtered and both spatially and temporally analyzed using discrete Fourier techniques. At all speeds tested for both designs, stall inception was characterized by growth of a small amplitude' rotating wave. The waves did not grow significantly until just prior to the instability, when exponential growth into fully-developed rotating stall occurred very rapidly, within 6-10 rotor revolutions. The amount of time the rotating waves could be detected prior to stall varied considerably with compressor operating condition and was largely dependent on the local slope of the compressor speedline characteristics. Stall warning times ranged from less than one-tenth of a second to more than two seconds for the same machine operated at different high speeds (above 60% design speed). The influence of compressibility effects are also discussed.