Browsing by Author "Moore, John"
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- 3-D flow and performance of a rocket pump inducer at design and off-design flow ratesDoan, Andrew W. (Virginia Tech, 1994-08-05)The ADP rocket pump inducer was studied computationally using a 3-D Navier-Stokes solver, The Moore Elliptic Flow Program. Design and off-design flow rates were simulated to qualitatively and quantitatively study the effects of flow rate on the flow and performance. Several views of the results were created to aid flow visualization. The 3-D laser measurements made by Rocketdyne were used for comparison. The velocity magnitudes as well as the flow patterns within the inducer match well between the calculated and measured results. The axial velocity distribution and the rotary stagnation pressure, losses, are predicted very well by the calculation. The internal flow patterns developed in the simulation as expected, with radial outflow in the blade boundary layers. The tip leakage flow formed a recirculation region, a toroidal shaped vortex at the tip leading edge of the blades. The associated backflow forms a blockage that varies with flow rate. The thermodynamic performance was evaluated by calculating the contributions to pressure rise, the pump characteristic, the contributions to moment of momentum, and the efficiency. The centrifugal effect and relative velocity effect were found to vary with flow rate. The effective inlet throat radius, which governs these two effects, changes with flow rate because of the recirculation blockage. The shear on the blades was found to produce a small fraction of the work in the inducer, and most was produced by the pressure difference across the blade. The inducer efficiency was about 89%, and increased with decreasing flow rate in the range of flow rates considered, from 89% to 110% of the design flow rate.
- 3-D flow and performance of a tandem-bladed rocket pump inducerExcoffon, Tony (Virginia Tech, 1992-04-05)This thesis presents the results of a three-dimensional flow calculation with a model of turbulent viscosity for a tandem-bladed inducer in air. The purpose is to understand the 3D flow development through the two blade rows and to compare the results of the calculation 'with experimental data. A literature review tells the story of the inducer from the flat-plate design to the tandem-bladed configuration and explains its role in cavitation management. The results of a previous 3D-calculation on the first blade row alone are summarized and the MEFP code is briefly described. The generation of a grid for the second blade row is presented in detail. Then, it is shown how this new grid is linked to the previous grid for the first blade row to get an overall calculation grid for the whole inducer. Two 2D blade-to-blade calculations are shown. They give an insight into the flow behavior through the inducer and allow a test of the grid. The results of the 3D-calculation are discussed and presented extensively with the velocity vectors, the static pressure contours and the rotary stagnation pressure contours on blade-to-blade, meridional and iso-8 vie"rs. The three passages of the second blade row appear to behave differently with respect to their position relative to the wake of the first blade row. The experimental data are used for comparison at three measurement planes in terms of pressure and velocity. They show a fairly good agreement. The three-dimensional calculation predicts also very well the work done and the efficiency of the overall inducer.
- Analysis of tilting-pad oil seals for high pressure centrifugal compressorsSalem, Khlifi (Virginia Tech, 1988-12-05)Oil ring seals are one major source of instability in high pressure centrifugal compressors. This thesis presents a method for analysis of an improved seal concept that has been used in very high pressure designs (900 PSI). The improved design uses a combination of ring seals and tilting pad bearing elements. The stable tilting pad is used to center the heavily grooved seal element. The eight stiffness and damping coefficients which represent the hydrodynamic forces between the journal and the seal assembly are computed by an automated computer code for evaluation of both the standard ring seal and the tilting pad elements. Both synchronous and nonsynchronous steady state characteristics have been included in the analysis. The nonsynchronous whirl of the rotor and its effects on the stiffness and damping coefficients of a 5 tilting pad seal have been given in the form of design curves. The effect of pad inertia which has been neglected in many bearing analysis codes has been incorporated in this seal analysis, and allowed the determination of the exact cross coupling stiffness and damping coefficients.
- CFD analysis and redesign of centrifugal impeller flows for rocket pumpsLupi, Alessandro (Virginia Tech, 1993-12-05)The analysis and redesign of a centrifugal impeller for a rocket pump is presented in this thesis. A baseline impeller was designed by Rocketdyne for the NASA Marshall Pump Consortium. Initially, the objective was to reduce the circumferential exit flow distortion of the baseline impeller. Later in the study, the objective became raising the head coefficient of the impeller. The study presented in this thesis was also undertaken to demonstrate current CFD capabilities for impeller design. A literature review includes an overview of centrifugal impeller geometries and configurations. Centrifugal impeller performance and secondary flows are discussed, and a summary of studies on the effects of impeller exit and diffuser inlet velocity distortion on diffuser performance is also presented. The flow calculation details and the results of the baseline impeller flow calculations are described. Fourteen redesigned impeller geometries were analyzed using the Moore Elliptic Flow Program, and the results were compared to the baseline geometry in terms of head rise, losses, and exit flow distortions. A final geometry was chosen; this geometry will be built and tested by Rocketdyne. The results show that backward blade lean can be effective in red using the exit flow distortion of the impeller. Tip slots or holes were not beneficial because of the large inlet boundary layer. Also, it appears possible to raise the head coefficient of the baseline impeller without creating excessive flow distortion. The planned testing is necessary to verify the predictions of the flow code.
- Computational study of 3D turbulent air flow in a helical rocket pump inducerLe Fur, Thierry (Virginia Tech, 1989-12-05)A computational study of the air flow in a helical rocket pump inducer has been performed using a 3-D elliptic flow procedure including viscous effects. The inlet flow is considered turbulent and fully developed. The geometric, definition of the inducer blade shape and the calculation grid are first presented, followed by a discussion of the flow calculation results displayed in various new graphical representations. The general characteristics expected from previous experimental and analytical work appear in the simulation and were quantitatively studied. The tip leakage flow observed has velocities of the order of the blade tip speed and is partially convected across the entire passage. The important boundary layer development on the blade pressure side and suction side creates radial outward flows, whereas a radial inward motion develops in the core region, with velocities of same order, and from shroud to hub. Secondary and tip leakage flows combine to give a region of high flow losses and blockage near the shroud wall, and the secondary flow pattern is nearly fully developed by the inducer exit. Original details were also resolved in the flow calculation. A circumferential vortex develops near the shroud, immediately upstream of the suction side of the swept-back leading edge. A simplified air-LH2 analogy permitted the prediction of cavitation inception in the liquid hydrogen pump, and the results obtained correspond qualitatively well with water flow visualizations. The accordance of the model with available air test data at the inlet and exit of the inducer is generally very good, with the total pressure losses in excellent agreement.
- Computational study of hub corner stall in an axial compressor rotorGailliot, John A. (Virginia Tech, 1995-08-15)The Deverson rotor, a single stage axial compressor designed to simulate a multistage axial compressor, was studied computationally using a 3-D Navier-Stokes solver, the Moore Elliptic Flow Program. A one equation, q-L, transitional turbulence model was used with MEFP for closure of the transport equations. The calculation was used to study the physics and flow mechanisms affecting hub corner stall. Preprocessing and post processing programs were written to aid this study, a grid generation program and a streakline visualization program, respectively. First, computational 2-D cascade studies were performed to study the effects of free stream turbulence level and incidence angle on suction surface boundary layer development. The results showed the correct trends in boundary layer transition and separation, loss production, and deviation angles. Velocity measurements taken at the exit of the Deverson rotor were made available by Rolls-Royce for comparison with the 3-D calculation results. The q-L turbulence model predicted the existence of the hub comer stall, but under predicted the size of the corner stall. It failed to predict the radial migration of the associated loss core. However, the calculation did reveal details of the flow that affect comer stall. These included boundary layer transition and separation on the suction surface, hub and suction surface secondary flows, and radial relief. Streaklines were useful in visualizing and understanding these flow details. A preliminary 3-D calculation was performed with a two-equation, q-w, turbulence model. This turbulence model more accurately predicted the comer stall including radial migration of the loss core.
- A computational study of the 3D flow and performance of a vaned radial diffuserAkseraylian, Dikran (Virginia Tech, 1996-08-14)A computational study was performed on a vaned radial diffuser using the MEFP (The Moore Elliptic Flow Program) flow code. The vaned diffuser studied by Dalbert et al. was chosen as a test case for this thesis. The geometry and inlet conditions were established from this study. The performance of the computational diffuser was compared to the test case diffuser. The CFD analysis was able to demonstrate the 3D flow within the diffuser. An inlet conditions analysis was performed to establish the boundary conditions at the diffuser inlet. The given inlet flow angles were reduced in order to match the specified mass flow rate. The inlet static pressure was held constant over the height of the diffuser. The diffuser was broken down into its subcomponents to study the effects of each component on the overall performance of the diffuser. The diffuser inlet region, which comprises the vaneless and semi-vaneless spaces, contains the greatest losses, 56%, but the highest static pressure rise, 54%. The performance at the throat was also evaluated and the blockage and pressure recovery were calculated. The results show the static pressure comparison for the computational study and the test case. The overall pressure rise of the computational study was in good agreement with the measured pressure rise. The static pressure and total pressure loss distributions in the inlet region, at the throat, and in the exit region of the diffuser were also analyzed. The flow development was presented for the entire diffuser. The 3D flow calculations were able to illustrate a leading edge recirculation at the hub, caused by an inlet skew and high losses at the hub, and the secondary flows in the diffuser convected the high losses. The study presented in this thesis demonstrated the flow development in a vaned diffuser and its subcomponents. The performance was evaluated by calculating the static pressure rise, total pressure losses, and throat blockage. It also demonstrated current CFD capabilities for diffusers using steady 3D flow analysis.
- Development of a new shock capturing formula for pressure correction methodsGupta, Ajay K. (Virginia Tech, 1993-12-05)Several methods have been developed to capture shock waves in turbo machinery flows, such as Moore's pressure correction procedure and Denton's time marching procedure. The time marching procedure is traditionally used for transonic flow calculations, whereas the pressure correction method is better suited for incompressible and subsonic flows. However, the focus of this research is on the Moore pressure correction flow code, the Moore Elliptical Flow Program (MEFP) , to calculate shock waves in transonic compressor fans. A new pressure interpolation method, the 2M formula, is developed to improve the shock capturing capabilities of the MEFP flow code. The 2M formula is a two Mach number dependent formula, with Mach numbers Mi and M i + 1. The previously used pressure interpolation method, the M&M formula, is a one Mach number dependent formula, using the maximum of Mi and Mi + 1 . In the development of the 2M formula, J.G. Moore's stability criterion is applied to the pressure correction equation such that the center point coefficient is greater than the sum of the other positive coefficients.
- The effect of boundary layer blowing in the corner region of a linear compressor cascade wind tunnelJames, Ralph William (Virginia Tech, 1995-06-05)A fundamental investigation of the flow in the endwall corner region of a linear compressor cascade wind tunnel and the effect of boundary layer blowing in this region was conducted using blade surface pressure tap measurements and five - hole prism probe measurements taken downstream of the cascade. The results are presented as a series of velocity vector plots, loss contour plots, and pitchwise mass - averaged loss coefficient plots. The angle of attack test range was from 5 to 29 degrees. For the corner region boundary layer blowing investigations, two slots were machined into the ends of a set of cascade blades, and an external air source was used as the blade slot jet air source. Tests were done for 19, 21, and 23 degrees angle of attack. The main effect of corner boundary layer blowing was a significant reduction in total pressure losses in the region along the blade span between the exterior portion of the corner boundary layer flow and the blade profile boundary layer flow.
- The effect of solidity on the pre- and post-stall flow in a linear compressor cascadeAinslie, Walter E. (Virginia Tech, 1988-05-02)An experimental investigation of the performance characteristics of a solid wall linear compressor cascade was conducted. The purpose of the experiments was to determine the effects of the blade row configuration parameters stagger and solidity on the pre-and post-stall behavior of the flow in the cascade. Tests were conducted at a solidity of 1.5, and for two stagger angles, 36.4 degrees and 25 degrees. The investigation included the use of high speed motion pictures with smoke flow visualization in the cascade, measurements of the total pressure and velocity of the flow upstream and downstream of the cascade, and measurements of the blade surface pressures. The experiments were conducted for a range of angle of attack from 0 degrees to 45 degrees. To determine the effects of solidity on the pre- and post-stall behavior of the flow in the cascade, the results obtained for the present 1.5 solidity cascade were compared to previous results from the same cascade tested at a solidity of 1.0. The flow in the two cascades was observed to be similar in nature, but the influence of the reduced blade loading in the high solidity cascade was apparent. For the higher solidity cascade, flow losses at low angle of attack were found to be larger, but stalling behavior was delayed.
- Effects of multiple incident shock waves on the flow in a transonic turbine cascadeDoughty, Roger L. (Virginia Tech, 1994-12-15)Turbine aerodynamic designers are currently focusing on unsteady passage flow to increase turbine performance. In particular, for high pressure turbine stages the effects of wakes and shocks shed from an upstream blade row on the downstream blade row need to be understood. Also, experimental data is needed for comparison with unsteady three-dimensional turbine stage calculations. Previous simulations of the unsteady shock/wake inlet flow field for a turbine rotor or stator used a rotating disk with radial bars upstream of a linear cascade. An alternate method of shock generation is developed here using a capped shock tube with multiple outlets to get a traveling system of three shock waves. Different lengths of tubing are used to get time delays between the shocks, which are then introduced at the top of a linear cascade of turbine blades and travel downwards (tangentially) along the leading edge. Advantages of this method include the absence of wakes and excellent two-dimensionality of the inlet shock waves. The period of the incoming shocks is easily adjustable to simulate different Strouhal numbers. Unsteady measurements of upstream total pressure, blade static pressures, and uncorrected downstream total pressure are made for a transonic mean flow with introduction of traveling shocks at M=1.3. An analytical solution (Bach and Lee, 1970) for the decay of cylindrical shock waves is used to estimate the behavior of flow variables other than pressure at the cascade inlet. The unsteady total pressure loss of the blade passage and the unsteady blade forces are measured with one shock passing and with three shocks passing at periods of 0.055 and 0.200 milliseconds. Loss is estimated as the normalized difference in unsteady total pressures and blade forces are integrated from seventeen unsteady surface pressure measurements. The Strouhal number for the 0.200 msec case is 2.9, which is typical of a high-pressure turbine nozzle or rotor. Periodic behavior in blade force and loss are observed for this case. Blade lift shows peak-to-peak variation of 6% and the estimated loss fluctuates by 100%. No change is observed in the average level of loss due to the incident shock waves.
- Effects of temperature transients on the stall and stall recovery aerodynamics of a multi-stage axial flow compressorDiPietro, 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.
- Experimental and numerical investigations of the off-design flow physics in a supersonic through-flow fan cascadeAndrew, Philip L. (Virginia Tech, 1992-08-01)The turbulent character of the supersonic wake of a linear cascade of fan airfoils has been studied experimentally using a two-component Laser Doppler Anemometer. The cascade was tested in the Virginia Polytechnic Institute and State University intermittent wind tunnel facility, where the experimental Mach and Reynolds numbers were 2.36 and 4.8 x 10⁶, respectively. In addition to mean flow measurements, Reynolds normal and shear stresses were measured as functions of cascade incidence angle and streamwise location in the near-wake and the far-wake. The extremities of profiles of both the mean and turbulent wake properties were found to be strongly influenced. by upstream shock-boundary-layer-interactions, the strength of which varied with cascade incidence. In contrast, the peak levels of turbulence properties within the shear layer were found to be largely independent of cascade incidence, and could be characterized in terms of the streamwise position only. This fact permitted the determination of the decay of the Reynolds shear stress, the production rate of turbulent kinetic energy, and the turbulent kinetic energy itself with streamwise location.
- Experimental evaluation of effective friction coefficient for liquid ring sealsDorton, David W. (Virginia Tech, 1991-08-05)Rotor dynamic analysis of liquid ring seals depends upon the correct specification of seal dynamic stiffness and damping characteristics. These are in turn dependent upon several parameters, including the friction holding force between the sealing face and the mating retaining ring. Designers currently assume a value for effective friction coefficient in order to utilize methods for prediction of response and stability. This thesis presents the results of testing on twelve actual seal rings of varying configuration at pressures of 689, 1378, 2068, and 2757 kPa in a static seal test rig to experimentally determine values of effective friction coefficient. The results are presented in graphical form as average effective friction coefficient versus eccentricity ratio for forward and backward motion of the rings.
- An experimental examination of the influence of trailing-edge coolant ejection on blade losses in transonic turbine cascadesBertsch, Remi (Virginia Tech, 1990-12-07)This thesis summarizes the results of an experimental study on transonic turbine blades in the presence of ejection of coolant in the direction of the flow from slots near the trailing edge. I t presents the effect of the trailing edge coolant ejection on the turbine blade aerodynamic efficiency.¹ The objective of this work is to contribute to the design of new turbine blades by giving loss data for cooled blades. Data were taken in the Virginia Polytechnic Institute & State University wind tunnel, which includes a two-dimensional transonic turbine cascade. The tunnel simulates supersonic discharge flows of turbine rotor blading in a linear cascade with trailing edges designed for ejection of cooling flow. Two blade designs, named Baseline and ULTRE, were tested. Experiments were performed on a transonic turbine cascade designed for a deflection of approximately 68 degrees and outlet Mach number of 1.14 for the Baseline blade and 1.2 for the ULTRE blade. Tests were carried out with CO₂ as coolant in order to ensure the proper simulation of the density ratio between coolant flow and main flow. Data were obtained for both the Baseline and ULTRE cascades with a good periodicity. The content of this thesis is limited to the aerodynamic aspects of coolant ejection. Heat transfer aspects are mentioned but not developed. The first part of this thesis reports on the theoretical considerations necessary for the understanding of the work done and describes the arrangement, instrumentation, and data acquisition system of the wind tunneL The second part of the thesis presents experimental results from tests carried out on both Baseline and ULTRE blades. The cascade tests cover an exit isentropic Mach number range of M2,it = 0.72 to 1.34 and four different ejection rates. 1 The efficiency being characterized by the total pressure loss in this work
- An experimental investigation of the turbulent flow in a closed compound channelKouroussis, Dimitrios (Virginia Tech, 1996-02-05)A three-component laser Doppler anemometer was used to measure the fully developed, turbulent flow in a closed, symmetric, smooth-wall compound channel. Measurements were made across one quadrant of the cross-section since the flow was assumed symmetric. Measurements were made for a single channel Reynolds number. All mean velocity components were calculated and are reported. The mean velocity field results are in good agreement with results reported for similar geometries. The vector plots and the axial vorticity distribution reveal the existence of secondary flow cells in both the main channel and the flood plain. The maximum values of the secondary velocities are at the comer region, on the interface between the main channel and the flood plain. In this region the mean velocity gradients are large, indicating that this might be an area of high turbulence production. The distributions of all Reynolds stresses across the cross-section are reported. The Reynolds stress distributions show peak values near the interface corner region and small values near the center-line and on the axes of symn1etry of the channel. The turbulence kinetic energy distribution verifies the existence of high turbulence energy fluid in the comer region.
- An experimental study of lifting and moving forces in air conveying systemsChardon, Sylvaine (Virginia Tech, 1992-02-04)An air conveying system uses pressurized air as a propelling force to lift and move articles. It is supplied by a fan into a plenum with a top surface that is a flat perforated plate. Air escapes through the openings, creating a layer that supports and drives the articles along. This thesis provides information on the lifting and moving forces. It summarizes the results of both analytical and experimental studies. Most of the effort is focused on an experimental procedure for measuring the actual forces on the objects being conveyed and data are used to verify the analytical models. The experiments are limited to straight holes and louvers located under the bottom of aluminum concave-bottom cans. In some tests, a flat disc has been fixed to the bottom of the cans. Measurements are made of the can motion on an actual section of conveyor.
- Extension of the finite volume method to laminar and turbulent flowNicholson, Stephen (Virginia Polytechnic Institute and State University, 1986)A method has been developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier Stokes equations. The entire calculation is performed in the physical domain. The method is currently limited to the calculation of attached flows. The features of the current method can be summarized as follows. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is not needed. Different time steps are used in the different governing equations to improve the convergence speed of the viscous calculations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes (length/height ≅ 1000). A special discretization technique is also used to stabilize these calculations which use long and thin control volumes. A special formulation of the energy equation is used to provide improved transient behavior of solutions which use the full energy equation. The method is then compared with a wide variety of test cases. The freestream Mach numbers range from 0.075 to 2.8 in the calculations. Transonic viscous flow in a converging diverging nozzle is calculated with the method; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good. Essentially incompressible turbulent boundary layer flow in an adverse pressure gradient is calculated and the computed distribution of mean velocity and shear stress are in good agreement with the measurements. At the other end of the Mach number range, a flat plate turbulent boundary layer with a freestream Mach number of 2.8 is calculated using the full energy equation; the computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.
- Flow losses in supersonic compressor cascadesBloch, Gregory S. (Virginia Tech, 1996-07-06)Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. The physical mechanisms that affect the flow in supersonic compressor cascades have been reviewed, including the changes in shock geometry that will occur with back pressure for both started and unstarted operation. Compressible fluid mechanics has been applied to the known shock geometry to obtain a physics-based engineering shock loss model that is applicable over the entire supersonic operating range of the cascade. Predictions from the present method have been compared to measurements and Navier-Stokes analyses of the L030-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. Son1e of the started comparisons exhibited good agreement, while others did not. A clear improvement has been demonstrated over previously published shock loss models, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss. The present shock loss model has been used as the basis of an overall loss prediction method by adding a constant value, representative of the viscous profile loss, to the predicted shock loss characteristics. The overall loss characteristics obtained in this manner showed good agreement with the experimental values over the most useful operating range of the cascade.
- Fluid flow and heat transfer in transonic turbine cascadesJanakiraman, S. V. (Virginia Tech, 1993-05-05)The aerodynamic and thermodynamic performance of an aircraft gas turbine directly affects the fuel consumption of the engine and the life of the turbine components. Hence, it is important to be able to understand and predict the fluid flow and heat transfer in turbine blades to enable the modifications and improvements in the design process. The use of numerical experiments for the above purposes is becoming increasingly common. The present thesis is involved with the development of a flow solver for turbine flow and heat transfer computations. A 3-D Navier-Stokes code, the Moore Elliptic Flow Program (MEFP) is used to calculate steady flow and heat transfer in turbine rotor cascades. Successful calculations were performed on two different rotor profiles using a one-equation q-L transitional turbulence model. A series of programs was developed for the post-processing of the output from the flow solver. The calculations revealed details of the flow including boundary layer development, trailing edge shocks, flow transition and stagnation and peak heat transfer rates. The calculated pressure distributions, losses, transition ranges, boundary layer parameters and peak heat transfer rates to the blade are compared with the available experimental data. The comparisons indicate that the q-L transitional turbulence model is successful in predicting flows in transonic turbine blade rows. The results also indicate that the calculated loss levels are independent of the gridding used while the heat transfer rate predictions improve with finer grids.