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Hah, Chunill

Publications and source records attributed to Hah, Chunill.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Toward Optimum Configuration of Circumferential Groove Casing Treatment in Transonic Compressor Rotors

The current paper first reviews experimental and numerical investigations to understand flow physics and to develop optimum configurations of circumferential grooves in compressor rotors. Circumferential grooves are used mainly to increase stall margin in axial compressors with small decrease in aerodynamic efficiency. Although circumferential groove casing treatment has been used widely, flow mechanisms of the circumferential grooves at near stall conditions are not well understood yet. Detailed time-dependent flow measurement inside tip gap in a high speed compressor is still a big challenge even though significant advance has been made in non-intrusive flow measurement technique. Therefore numerical approaches have been used to study relevant flow physics. However, optimum design of circumferential grooves to a given compressor with the computational tools is not practical yet. In the present paper, various investigations to study flow physics of circumferential groove casing treatment in axial compressor are reviewed first. Possible missing flow physics are identified and future research efforts for the optimum design are discussed.

Hah, Chunill↗
NASA NTRS2011

Reynolds-Averaged Navier-Stokes Solutions to Flat Plate Film Cooling Scenarios

The predictions of several Reynolds-Averaged Navier-Stokes solutions for a baseline film cooling geometry are analyzed and compared with experimental data. The Fluent finite volume code was used to perform the computations with the realizable k-epsilon turbulence model. The film hole was angled at 35 to the crossflow with a Reynolds number of 17,400. Multiple length-to-diameter ratios (1.75 and 3.5) as well as momentum flux ratios (0.125 and 0.5) were simulated with various domains, boundary conditions, and grid refinements. The coolant to mainstream density ratio was maintained at 2.0 for all scenarios. Computational domain and boundary condition variations show the ability to reduce the computational cost as compared to previous studies. A number of grid refinement and coarsening variations are compared for further insights into the reduction of computational cost. Liberal refinement in the near hole region is valuable, especially for higher momentum jets that tend to lift-off and create a recirculating flow. A lack of proper refinement in the near hole region can severely diminish the accuracy of the solution, even in the far region. The effects of momentum ratio and hole length-to-diameter ratio are also discussed.

Johnson, Perry L.↗
NASA NTRS2010

Near Stall Flow Analysis in the Transonic Fan of the RTA Propulsion System

Turbine-based propulsion systems for access to space have been investigated at NASA Glenn Research center. A ground demonstrator engine for validation testing has been developed as a part of the program. The demonstrator, the Revolutionary Turbine Accelerator (RTA-1), is a variable cycle turbofan ramjet designed to transition from an augmented turbofan to a ramjet that produces the thrust required to accelerate the vehicle to Mach 4. The RTA-1 is designed to accommodate a large variation in bypass ratio from sea level static to Mach 4 flight condition. A key component of this engine is a new fan stage that accommodates these large variations in bypass ratio and flow ranges. In the present study, unsteady flow behavior in the fan of the RTA-1 is studied in detail with large eddy simulation (LES) and the numerical results are compared with measured data. During the experimental study of the fan stage, humming sound was detected at 100 % speed near stall operation. The main purpose of the study is to investigate details of the unsteady flow behavior at near stall operation and to identify a possible cause of the hum. The large eddy simulation of the current flow field reproduces main features of the measured flow very well. The LES simulation indicates that non-synchronous flow instability develops as the fan operates toward the stall limit. The FFT analysis of the calculated wall pressure shows that the rotating flow instability has the characteristic frequency that is about 50% of the blade passing frequency.

Hah, Chunill↗
NASA NTRS2009

Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements

In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.

Hah, Chunill↗
NASA NTRS2009

Large Eddy Simulation of Transonic Flow Field in NASA Rotor 37

The current paper reports on numerical investigations on the flow characteristics in a transonic axial compressor, NASA Rotor 37. The flow field was used previously as a CFD blind test case conducted by American Society of Mechanical Engineers in 1994. Since the CFD blind-test exercise, many numerical studies on the flow field in the NASA Rotor 37 have been reported. Although steady improvements have been reported in both numerical procedure and turbulence closure, it is believed that all the important aspects of the flow field have not been fully explained with numerical studies based on the Reynolds Averaged Navier-Stokes (RANS) solution. Experimental data show large dip in total pressure distribution near the hub at downstream of the rotor at 100% rotor speed. Most original numerical solutions from the blind test exercise did not predict this total pressure deficit correctly. This total pressure deficit at the rotor exit was attributed to a hub corner flow separation by the author. Several subsequent numerical studies with different turbulence closure model also calculated this dip in total pressure rise. Also, several studies attributed this total pressure deficit to a small leakage flow coming from the hub in the test article. As the experimental study cannot be repeated, either explanation cannot be validated. The primary purpose of the current investigation is to investigate the transonic flow field with both RANS and a Large Eddy Simulation (LES). The RANS approach gives similar results presented at the original blind test exercise. Although the RANS calculates higher overall total pressure rise, the total pressure deficit near the hub is calculated correctly. The numerical solution shows that the total pressure deficit is due to a hub corner flow separation. The calculated pressure rise from the LES agrees better with the measured total pressure rise especially near the casing area where the passage shock interacts with the tip clearance vortex and flow becomes unsteady due to this interaction. The LES simulation also calculates the total pressure rise deficit near the hub and it agrees well with the measured data.

Hah, Chunill↗
NASA NTRS2007

Aerodynamic Inner Workings of Circumferential Grooves in a Transonic Axial Compressor

The current paper reports on investigations of the fundamental flow mechanisms of circumferential grooves applied to a transonic axial compressor. Experimental results show that the compressor stall margin is significantly improved with the current set of circumferential grooves. The primary focus of the current investigation is to advance understanding of basic flow mechanics behind the observed improvement of stall margin. Experimental data and numerical simulations of a circumferential groove were analyzed in detail to unlock the inner workings of the circumferential grooves in the current transonic compressor rotor. A short length scale stall inception occurs when a large flow blockage is built on the pressure side of the blade near the leading edge and incoming flow spills over to the adjacent blade passage due to this blockage. The current study reveals that a large portion of this blockage is created by the tip clearance flow originating from 20% to 50% chord of the blade from the leading edge. Tip clearance flows originating from the leading edge up to 20% chord form a tip clearance core vortex and this tip clearance core vortex travels radially inward. The tip clearance flows originating from 20% to 50% chord travels over this tip clearance core vortex and reaches to the pressure side. This part of tip clearance flow is of low momentum as it is coming from the casing boundary layer and the blade suction surface boundary layer. The circumferential grooves disturb this part of the tip clearance flow close to the casing. Consequently the buildup of the induced vortex and the blockage near the pressure side of the passage is reduced. This is the main mechanism of the circumferential grooves that delays the formation of blockage near the pressure side of the passage and delays the onset of short length scale stall inception. The primary effect of the circumferential grooves is preventing local blockage near the pressure side of the blade leading edge that directly determines flow spillage around the leading edge. The circumferential grooves do not necessarily reduce the over all blockage built up at the rotor tip section.

Hah, Chunill↗
NASA NTRS2006

Unsteady Analysis of Turbine Main Flow Coupled with Secondary Air Flow

Two numerical approaches are used to model the interaction between the turbine main gas flow and the wheelspace cavity seal flow. The 3-D, unsteady Reynolds-averaged Navier-Stokes equations are solved with a CFD code based on a structured grid to study the interaction between the turbine main gas flow and the wheelspace cavity seal flow. A CFD code based on an unstructured grid is used to solve detailed flow feature in the cavity seal which has a complex geometry. The numerical results confirm various observations from earlier experimental studies under similar flow conditions. When the flow rate through the rim cavity seal is increased, the ingestion of the main turbine flow into the rim seal area decreases drastically. However, a small amount of main gas flow is ingested to the rim seal area even with very high level of seal flow rate. This is due to the complex nature of 3-D, unsteady flow interaction near the hub of the turbine stage.

Hah, Chunill↗
NASA NTRS2006

Fabrication Materials for a Closed Cycle Brayton Turbine Wheel

A multidisciplinary analysis of a radial inflow turbine rotor is presented. This work couples high-fidelity fluid, structural, and thermal simulations in a seamless multidisciplinary analysis to investigate the consequences of material selection. This analysis extends multidisciplinary techniques previously demonstrated on rocket turbopumps and hypersonic engines. Since no design information is available for the anticipated Brayton rotating machinery, an existing rotor design (the Brayton Rotating Unit (BRU)) was used in the analysis. Steady state analysis results of a notional turbine rotor indicate that stress levels are easily manageable at the turbine inlet temperature, and stress levels anticipated using either superalloys or ceramics.

Khandelwal, Suresh↗
NASA NTRS2005

Satellite Propellant Pump Research

NASA Glenn initiated a satellite propellant pump technology demonstration program. The goal was to demonstrate the technologies for a 60 percent efficient pump at 1 gpm flow rate and 500 psia pressure rise. The pump design and analysis used the in-house developed computer codes named PUMPA and HPUMP3D. The requirements lead to a 4-stage impeller type pump design with a tip diameter of 0.54 inches and a rotational speed of 57,000 rpm. Analyses indicated that flow cavitation was not a problem in the design. Since the flow was incompressible, the stages were identical. Only the 2-stage pump was designed, fabricated, assembled, and tested for demonstration. Water was selected as the surrogate fluid for hydrazine in this program. Complete mechanical design including stress and dynamic analyses were conducted. The pump was driven by an electric motor directly coupled to the impellers. Runs up to 57,000 rpm were conducted, where a pressure rise of 200 psia at a flow rate of 0.8 gpm was measured to validate the design effort.

Schneider, Steven J.↗
NASA NTRS2003

A Compliant Casing for Transonic Axial Compressors

A viewgraph presentation on the concept of compliant casing for transonic axial compressors is shown. The topics include: 1) Concept for compliant casing; 2) Rig and facility details; 3) Experimental results; and 4) Numerical results.

Bloch, Gregory S.↗
NASA NTRS2003

2002 Computing and Interdisciplinary Systems Office Review and Planning Meeting

The technologies necessary to enable detailed numerical simulations of complete propulsion systems are being developed at the NASA Glenn Research Center in cooperation with NASA Glenn s Propulsion program, NASA Ames, industry, academia and other government agencies. Large scale, detailed simulations will be of great value to the nation because they eliminate some of the costly testing required to develop and certify advanced propulsion systems. In addition, time and cost savings will be achieved by enabling design details to be evaluated early in the development process before a commitment is made to a specific design. This year s review meeting describes the current status of the NPSS and the Object Oriented Development Kit with specific emphasis on the progress made over the past year on air breathing propulsion applications for aeronautics and space transportation applications. Major accomplishments include the first 3-D simulation of the primary flow path of a large turbofan engine in less than 15 hours, and the formal release of the NPSS Version 1.5 that includes elements of rocket engine systems and a visual based syntax layer. NPSS and the Development Kit are managed by the Computing and Interdisciplinary Systems Office (CISO) at the NASA Glenn Research Center and financially supported in fiscal year 2002 by the Computing, Networking and Information Systems (CNIS) project managed at NASA Ames, the Glenn Aerospace Propulsion and Power Program and the Advanced Space Transportation Program.

Lytle, John↗
NASA NTRS2003

Mechanisms of Rotating Instability in Axial Compressors Investigated

Rotating instability is a phenomenon that occurs in the tip flow region of axial compressor stages during stable operation. It can be observed in highly staggered rotors with significant tip clearance and is strongest at high-load operating points where the characteristic levels off. In this condition, the single-stage fan under investigation radiates an audible, whistling tone, and wall pressure spectra in the vicinity of the rotor disk exhibit nonrotational components. The graph shows the spectrum of static pressure at a point on the endwall near the leading edge. A hump appears at roughly half of the blade passing frequency (BPF) and is characteristic of rotating instability. A computational model was developed at the NASA Glenn Research Center to investigate the mechanism behind this phenomenon. A three-dimensional steady Navier-Stokes code that has been successfully tested for a wide range of turbomachinery flows was modified to execute a time-accurate simulation of the full annulus of the compressor. At the inlet of the computational domain, the total pressure, total temperature, and two velocity components are specified. Since no unsteady measurements of static pressure or other flow variables were available downstream of the rotor, circumferentially averaged static pressure was specified on the shroud at the outlet of the computational domain. A three-dimensional view of the vortex from the numerical model is shown. Particle traces released near the leading edge tip have rolled up to illustrate the tip clearance vortex. Flow near the trailing edge is pushed forward by the axially reversed flow. It then interacts with the tip clearance flow and the incoming flow and results in the rotating instability vortex, the core of which is illustrated by total pressure shading on planes located successively downstream. The rotating instability vortex is formed periodically midway between the blades and moves toward the pressure side of the passage. The unsteady behavior of this vortex structure is the main mechanism of the rotating instability is shown. The numerical model can be used to detect any possible occurrence of rotating instability when the tip clearance increases during engine service.

Hah, Chunill↗
NASA NTRS1993

Three-dimensional unstructured grid Euler method applied to turbine blades

Flow through a turbine annular cascade is calculated using a three-dimensional Euler method based on unstructured tetrahedral meshes. The equations are integrated in time using an explicit Runge-Kutta time-stepping scheme. The inviscid flux terms are discretized using a cell-centered finite-volume formulation with upwind flux-difference splitting. The tetrahedral meshes around the turbine blade are generated using an advancing-front technique with forced geometric periodicity between the blades. Good agreement is obtained between the present calculation and the experiment for both surface pressure distribution and flow behavior in the passage between the blades, demonstrating the capability of the present methodology for turbomachinery flow applications.

Kwon, Oh J.↗
NASA NTRS1993

Bgridi - Interactive three-dimensional turbomachinery grid generation system with applications

An interactive, graphics-based grid generation system for 3D turbomachinery geometries has been developed. The system consists of separate modules for geometry modeling and grid generation. The grid generation method generates a series of 2D grids in the blade-to-blade passage to build up the 3D grid. A Poisson equation with forcing functions selected to control orthogonality and spacing on all boundaries is solved to generate the grid. Interactive definition of Bezier curves and surfaces as internal boundaries is used to improve control of grid quality. A multi-block data structure simpifies the creation of structured H-grids about complex turbomachinery geometries. The multi-block strategy facilitates the creation of a grid in the tip flow region, which is an important contributor to losses within the passage. The grid generation system is applied to several complex geometries including a simple radial impeller, a conventional turbine rotor, the SSME LOX turbine rotor with and without a tip flow grid, a multi-passage inducer-impeller, and a tip-flow cascade. In addition, a calculation of the tip-flow cascade flow field is shown.

Shoemaker, J. M.↗
NASA NTRS1992

Three-dimensional flow phenomena in a transonic, high-through-flow, axial-flow compressor stage

A detailed aerodynamic study of a transonic, high-through-flow, single stage compressor is presented. The compressor stage was comprised of a low-aspect-ratio rotor combined alternately with two different stator designs. Both experimental and numerical studies are conducted to understand the details of the complex flow field present in this stage. Aerodynamic measurements using high-frequency, Kulite pressure transducers and conventional probes are compared with results from a three-dimensional viscous flow analysis. A steady multiple blade row approach is used in the numerical technique to examine the detailed flow structure inside the rotor and the stator passages. The comparisons indicate that many flow field features are correctly captured by viscous flow analysis, and therefore unmeasured phenomena can be studied with some level of confidence.

Copenhaver, William W.↗
NASA NTRS1992

A critical evaluation of a three-dimensional Navier-Stokes method as a tool to calculate transonic flows inside a low-aspect-ratio compressor

A numerical study to evaluate a three-dimensional Navier-Stokes method as a tool to predict the detailed flow field inside a low-aspect-ratio compressor at various operating conditions was conducted. The details of the flow structure inside a low aspect ratio compressor (three-dimensional shock structure, shock-boundary layer interaction, and tip leakage vortex) and the overall aerodynamic performance at design and off-design conditions are numerically analyzed and the results are compared with the available experimental data. The flow field inside a state-of-the-art transonic compressor is used for the purpose of the evaluation.

Hah, Chunill↗
NASA NTRS1991

A viscous flow study of shock-boundary layer interaction, radial transport, and wake development in a transonic compressor

A numerical study based on the 3D Reynolds-averaged Navier-Stokes equation has been conducted to investigate the detailed flow physics inside a transonic compressor. 3D shock structure, shock-boundary layer interaction, flow separation, radial mixing, and wake development are all investigated at design and off-design conditions. Experimental data based on laser anemometer measurements are used to assess the overall quality of the numerical solution. An additional experimental study to investigate end-wall flow with a hot-film was conducted, and these results are compared with the numerical results. Detailed comparison with experimental data indicates that the overall features of the 3D shock structure, the shock-boundary layer interaction, and the wake development are all calculated very well in the numerical solution. The numerical results are further analyzed to examine the radial mixing phenomena in the transonic compressor. A thin sheet of particles is injected in the numerical solution upstream of the compressor. The movement of particles is traced with a 3D plotting package. This numerical survey of tracer concentration reveals the fundamental mechanisms of radial transport in this transonic compressor.

Hah, Chunill↗