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Browsing by Author "Davis, Milton W. Jr."

Now showing 1 - 7 of 7
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    The Dynamics of Stall and Surge Behavior in Axial-Centrifugal Compressors
    Cousins, William T. (Virginia Tech, 1997-12-02)
    The phenomena of stall and surge in axial-centrifugal compressors is investigated through high-response measurements of both the pressure field and the flowfield throughout the surge cycle. A unique high-response forward-facing and aft-facing probe provides flow information. Several axial-centrifugal compressors are examined, both in compressor rigs and engines. Extensive discussion is presented on the differences in axial and centrifugal rotors and their effect on the system response characteristics. The loading parameters of both are examined and data is presented that shows the increased tolerance of the centrifugal stage to instability. The dynamics of the compressor blade response are shown to be related to the transport time of a fluid particle moving through a blade passage. The data presented provides new insight into the dynamic interactions that occur prior to and during stall and surge. In addition, the inception of rotating stall and the inception of surge are shown to be the same phenomena . An analytical dynamic model (DYNTECC) is applied to one of the compression systems and the results are compared to data. The results show that the model can capture the global effects of rotating stall and surge. The data presented, along with the analytical results, provide useful information for the design of active and passive stall control systems.
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    Effectiveness of a Serpentine Inlet Duct Flow Control Scheme at Design and Off-Design Simulated Flight Conditions
    Rabe, Angela C. (Virginia Tech, 2003-08-01)
    An experimental investigation was conducted in a static ground test facility to determine the flow quality of a serpentine inlet duct incorporating active flow control for several simulated flight conditions. The total pressure distortion at the aerodynamic interface plane (AIP) was then used to predict the resulting stability for a compression system. This study was conducted using a model of a compact, low observable, engine inlet duct developed by Lockheed Martin. A flow control technique using air injection through microjets at 1% of the inlet mass flow rate was developed by Lockheed Martin to improve the quality of the flow exiting the inlet duct. Both the inlet duct and the flow control technique were examined at cruise condition and off-design simulated flight conditions (angle of attack and asymmetric distortion). All of the experimental tests were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number of 1.76*105 based on the hydraulic diameter at the inlet throat. For each of the flight conditions tested, the flow control scheme was found to improve the flow uniformity and reduce the inlet distortion at the AIP. For simulated cruise condition, the total pressure recovery was improved by ~2% with the addition of flow control. For the off-design conditions of angle of attack and asymmetric distortion, the total pressure recovery was improved by 1.5% and 2% respectively. All flight conditions tested showed a reduction in circumferential distortion intensity with flow control. The cruise condition case showed reduced maximum circumferential distortion of 70% with the addition of flow control. A reduction in maximum circumferential distortion of 40% occurred for the angle of attack case with flow control, and 30% for the asymmetric distortion case with flow control. The inlet total pressure distortion was used to predict the changes in stability margin of a compression system due to design and off-design flight conditions and the improvement of the stability margin with the addition of flow control. A parallel compressor model (DYNTECC) was utilized to predict changes in the stability margin of a representative compression system (NASA Stage 35). Without flow control, all three cases show similar reduced stability margins on the order of 30% of the original stability margin for NASA Stage 35 at 70% corrected rotor speed. With the addition of flow control, the cruise condition tested improved the stability margin to 80% of the original value while the off-design conditions recover to 60% of the original margin. Overall, the flow control has been found to be extremely beneficial in improving the operating range of a compression system for the same inlet duct without flow control.
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    Effects of temperature transients on the stall and stall recovery aerodynamics of a multi-stage axial flow compressor
    DiPietro, Anthony Louis (Virginia Tech, 1997-02-05)
    An experimental investigation into the effects of inlet temperature transients on the stall and stall recovery aerodynamics of a low speed multi-stage axial flow compressor has been presented. Experiments were conducted on a low speed multi-stage axial flow compression system to demonstrate how a compressor dynamically stalls or recovers from a rotating stall operating condition during an inlet temperature transient. The specific effects of the inlet temperature transients on the compressor rotor blade flow physics during the dynamic stall or rotating stall recovery events of the axial flow compression system have been presented. In one experiment, a full recovery from a rotating stall operating condition was successfully accomplished on the low speed multistage axial flow compressor. Explanations for the axial flow compression system dynamic stall and rotating stall recovery processes during inlet temperature transients have been presented. The method utilized for inducing the rotating stall recovery on the compression system has been proposed as a possible new technique for active recovery from rotating stall for single and multi-stage axial flow compression systems.
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    An Improved Streamline Curvature Approach for Off-Design Analysis of Transonic Compression Systems
    Boyer, Keith M. (Virginia Tech, 2001-04-09)
    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. Overall and spanwise comparisons demonstrated that the improved model gives reasonable performance trends and generally accurate results, indicating that the physical understanding of the blade effects and the flow physics that underlie the loss model improvements are correct and realistic. The new model is unique in its treatment of shock losses, and is considered a significant improvement for fundamentally based, accurate throughflow numerical approximations. The specific SLC model used here is employed in a novel numerical approach — the Turbine Engine Analysis Compressor Code (TEACC). With implementation of the improved SLC model and additional recommendations presented within this report, the TEACC method offers increased potential for accurate analysis of complex, engine-inlet integration issues, such as time-variant inlet distortion.
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    Multidimensional dynamic compression system modeling
    Lindau, Jules Washington (Virginia Tech, 1995-06-05)
    A more robust method for solving the governing equations of a one-dimensional stage-by-stage dynamic compression system model was developed and validated. The improved method was then applied to two-dimensional post-stall models. The improvement in robustness was achieved by modeling the governing equations with upwind differencing and use of implicit time integration. A special form of upwind flux, flux difference splitting with source term treatment, FDSS, was developed for the model. A two-dimensional axisymmetric model was developed to allow post-stall modeling of split flowpath systems such as turbofans. This model was an entirely new concept. Additionally, a two-dimensional axial-circumferential model of rotating stall cell development and propagation was developed based on previous work. All of the models developed applied upwind differencing techniques to improve upon central-difference methods.
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    A stage-by-stage post-stall compression system modeling technique: methodology, validation, and application
    Davis, Milton W. Jr. (Virginia Polytechnic Institute and State University, 1986)
    A one-dimensional, stage-by-stage axial compression system mathematical model has been constructed which can describe system behavior during post-stall events such as surge and rotating stall. The model uses a numerical technique to solve the nonlinear conservation equations of mass, momentum, and energy. Inputs for blade forces and shaft work are provided by a set of quasi-steady stage characteristics modified by a first order lagging equation to simulate dynamic stage characteristics. The model was validated with experimental results for a three-stage, low-speed compressor and a nine-stage, high-pressure compressor. Using these models, a parametric study was conducted to determine the effect of inlet resistance, combustor performance, heat transfer, and stage characteristic changes due to hardware modification on post—stall system behavior.
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    A three-dimensional turbine engine analysis compressor code (TEACC) for steady-state inlet distortion
    Hale, Alan A. (Virginia Tech, 1996-12-15)
    Modem high-performance military aircraft are subjected to rapid flight maneuvers which place great operational demands on their compression system by producing highly distorted flow to the compressor. Inlet distortion generally reduces the engine compressor stability margin and may induce compressor surge at high rotational speeds, or rotating stall at lower rotational speeds. Therefore, a computational fluid dynamics (CFD) based compressor simulation would be very useful in the design, test, and analysis process since it gives additional information with inexpensive modifications. A new CFD simulation called the Turbine Engine Analysis Compressor Code (TEACC) was designed to meet these requirements. This code solves the compressible 3D Euler equations modified to include turbomachinery source terms which simulates the effect of the compressor blades. The source terms are calculated for each blade row by the application of a streamline curvature code. A methodology was developed for calculating turbomachinery source terms and distributing them axially, radially, and circumferentially while maintaining a sensitivity to strong inlet distortion. TEACC was compared with experimental data from NASA Rotor 1 B, a transonic rotor. Experimental data from Rotor 1 B were available for comparison with TEACC results for a clean inlet and for an inlet distortion produced by a 90-degree, one-per-revolution screen. TEACC results compared very well with experimental data with a clean inlet. Comparison with experimental data with inlet distortion demonstrated TEACC's ability to characterize the compressor overall, and to accurately predict the magnitude and shape of exit total temperature and exit total pressure in the distorted region. TEACC calculated the overall character of exit total pressure and exit total temperature in the nondistorted region, identifying the location of the largest value just after the inlet distortion and the decrease in exit total values through the nondistorted region in the direction of rotation.