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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Multi-airfoil Navier-Stokes simulations of turbine rotor-stator interaction

An accurate numerical analysis of the flows associated with rotor-stator configurations in turbomachinery can be extremely helpful in optimizing the performance of turbomachinery. In this study the unsteady, thin-layer, Navier-Stokes equations in two spatial dimensions are solved on a system of patched and overlaid grids for a rotor-stator configuration from an axial turbine. The governing equations are solved using a finite-difference, upwind algorithm that is set in an iterative, implicit framework. Results in the form of pressure contours, time-averaged pressures, unsteady pressure amplitudes and phase are presented. The numerical results are compared with experimental data and the agreement is found to be good. The results are also compared with those of an earlier study which used only one rotor and one stator. The current study uses multiple rotors and stators and a pitch ratio that is much closer to the experimental ratio. Consequently the results of this study are found to be closer to the experimental data.

Rai, Man Mohan↗
NASA NTRS1974

Variably positioned guide vanes for aerodynamic choking

A choking device to cause a sonic barrier to be formed which reduces the transmission of noise in a direction opposed to the direction of air flow in a compressor that may be part of an aircraft gas turbine engine is described. The noise reduction is accomplished by proper shaping and movement of inlet guide vanes, and an actuator is connected to selected guide vanes to effect movement by programmed amounts as required to choke or partially choke within the design range of the axial-flow-air compressor.

Chestnutt, D.↗
NASA NTRS1945

Data for Design of Entrance Vanes from Two-Dimensional Tests of Airfoils in Cascade

As a part of a program of the NACA directed toward increasing the efficiency of compressors and turbines, data were obtained for application to the design of entrance vanes for axfax-flow compressors or turbines. A series of blower-blade sections with relatively high critical speeds have been developed for turning air efficiently from 0 deg to 80 deg starting with an axial direction. Tests were made of five NACA 65-series blower blades (modified NACA 65(216)-010 airfoils) and of four experimentally designed blower blades in a stationary cascade at low Mach numbers. The turning effectiveness and the pressure distributions of these blade sections at various angles of attack were evaluated over a range of solidities near 1. Entrance-vane design charts are presented that give a blade section and angle of attack for any desired turning angle. The blades thus obtained operate with peak-free pressure distributions. Approximate critical Mach numbers were calculated from the pressure distributions.

FLOW-VANES, GUIDE↗
NASA NTRS2015

Design of an Object-Oriented Turbomachinery Analysis Code: Initial Results

Performance prediction of turbomachines is a significant part of aircraft propulsion design. In the conceptual design stage, there is an important need to quantify compressor and turbine aerodynamic performance and develop initial geometry parameters at the 2-D level prior to more extensive Computational Fluid Dynamics (CFD) analyses. The Object-oriented Turbomachinery Analysis Code (OTAC) is being developed to perform 2-D meridional flowthrough analysis of turbomachines using an implicit formulation of the governing equations to solve for the conditions at the exit of each blade row. OTAC is designed to perform meanline or streamline calculations; for streamline analyses simple radial equilibrium is used as a governing equation to solve for spanwise property variations. While the goal for OTAC is to allow simulation of physical effects and architectural features unavailable in other existing codes, it must first prove capable of performing calculations for conventional turbomachines.OTAC is being developed using the interpreted language features available in the Numerical Propulsion System Simulation (NPSS) code described by Claus et al (1991). Using the NPSS framework came with several distinct advantages, including access to the pre-existing NPSS thermodynamic property packages and the NPSS Newton-Raphson solver. The remaining objects necessary for OTAC were written in the NPSS framework interpreted language. These new objects form the core of OTAC and are the BladeRow, BladeSegment, TransitionSection, Expander, Reducer, and OTACstart Elements. The BladeRow and BladeSegment consumed the initial bulk of the development effort and required determining the equations applicable to flow through turbomachinery blade rows given specific assumptions about the nature of that flow. Once these objects were completed, OTAC was tested and found to agree with existing solutions from other codes; these tests included various meanline and streamline comparisons of axial compressors and turbines at design and off-design conditions.

compressors↗
NASA NTRS2015

Design of an Object-Oriented Turbomachinery Analysis Code: Initial Results

Performance prediction of turbomachines is a significant part of aircraft propulsion design. In the conceptual design stage, there is an important need to quantify compressor and turbine aerodynamic performance and develop initial geometry parameters at the 2-D level prior to more extensive Computational Fluid Dynamics (CFD) analyses. The Object-oriented Turbomachinery Analysis Code (OTAC) is being developed to perform 2-D meridional flowthrough analysis of turbomachines using an implicit formulation of the governing equations to solve for the conditions at the exit of each blade row. OTAC is designed to perform meanline or streamline calculations; for streamline analyses simple radial equilibrium is used as a governing equation to solve for spanwise property variations. While the goal for OTAC is to allow simulation of physical effects and architectural features unavailable in other existing codes, it must first prove capable of performing calculations for conventional turbomachines. OTAC is being developed using the interpreted language features available in the Numerical Propulsion System Simulation (NPSS) code described by Claus et al (1991). Using the NPSS framework came with several distinct advantages, including access to the pre-existing NPSS thermodynamic property packages and the NPSS Newton-Raphson solver. The remaining objects necessary for OTAC were written in the NPSS framework interpreted language. These new objects form the core of OTAC and are the BladeRow, BladeSegment, TransitionSection, Expander, Reducer, and OTACstart Elements. The BladeRow and BladeSegment consumed the initial bulk of the development effort and required determining the equations applicable to flow through turbomachinery blade rows given specific assumptions about the nature of that flow. Once these objects were completed, OTAC was tested and found to agree with existing solutions from other codes; these tests included various meanline and streamline comparisons of axial compressors and turbines at design and off-design conditions.

object-oriented programming↗
NASA NTRS1980

Aerodynamics of advanced axial-flow turbomachinery

A multi-task research program on aerodynamic problems in advanced axial-flow turbomachine configurations was carried out at Iowa State University. The elements of this program were intended to contribute directly to the improvement of compressor, fan, and turbine design methods. Experimental efforts in intra-passage flow pattern measurements, unsteady blade row interaction, and control of secondary flow are included, along with computational work on inviscid-viscous interaction blade passage flow techniques. This final report summarizes the results of this program and indicates directions which might be taken in following up these results in future work. In a separate task a study was made of existing turbomachinery research programs and facilities in universities located in the United States. Some potentially significant research topics are discussed which might be successfully attacked in the university atmosphere.

Serovy, G. K.↗
NASA NTRS1995

Enhanced Boiling-Metal Cooling Of Vanes Exposed To Hot Gases

Incorporation of automatic, self-powered jet pumps proposed to enhance boiling-liquid-metal cooling of vanes exposed to hot gases. In original intended application, vanes and probes thrust-vector-control devices inserted in supersonic flows of hot gases in rocket-engine nozzles; this cooling concept also applicable to vanes and blades in high-performance turbine engines. In further improvement, additional axial and transverse slots added to coolant passages in vane or probe and to coolant reservoir. Slots reduce stresses caused by thermal expansion and contraction of solid coolant.

Osofsky, I. B.↗
NASA NTRS2013

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence

Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to −51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and −36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg

Flegel-McVetta, Ashlie B.↗
NASA NTRS2007

Aspects of Transition in Turbomachines

This talk provides a description of several types of transition encountered in turbomachines. It is based largely on personal experience of the detection of transition in turbomachines. Examples are taken from axial compressors, axial turbines and radial turbines. The illustrations are concerned with transition in steady and unsteady boundary layers that develop under the influence of two-dimensional and three-dimensional flow fields.

Hodson, H. P.↗
NASA NTRS1956

Aerodynamic Design of Axial-flow Compressors. Volume III

Chapters XI to XIII concern the unsteady compressor operation arising when compressor blade elements stall. The fields of compressor stall and surge are reviewed in Chapters XI and XII, respectively. The part-speed operating problem in high-pressure-ratio multistage axial-flow compressors is analyzed in Chapter XIII. Chapter XIV summarizes design methods and theories that extend beyond the simplified two-dimensional approach used previously in the report. Chapter XV extends this three-dimensional treatment by summarizing the literature on secondary flows and boundary layer effects. Charts for determining the effects of errors in design parameters and experimental measurements on compressor performance are given in Chapters XVI. Chapter XVII reviews existing literature on compressor and turbine matching techniques.

COMPRESSORS - PRESSURE RATIO↗
DOE OSTI2024

Wind plant wake losses: Disconnect between turbine actuation and control of plant wakes with engineering wake models

Wake losses from neighboring plants may become a major factor in wind plant design and control as additional plants are constructed in areas with high wind resource availability. Because plant wakes span a large range of physical scales, from turbine rotor diameter to tens of kilometers, it is unclear whether conventional wake models or turbine control strategies are effective at the plant scale. Wake steering and axial induction control are evaluated in the current work as means of reducing the impact of neighboring wind plants on power and levelized cost of electricity. FLOw Redirection and Induction in Steady State (FLORIS) simulations were performed with the Gauss–Curl Hybrid and TurbOPark wake models as well as two operation and maintenance models to investigate control setpoint sensitivity to wake representation and economic factors. Both wake models estimate losses across a range of atmospheric conditions, although the wake loss magnitude is dependent on the wake model. Annual energy production and levelized cost of electricity are driven by wind direction frequency, with frequently aligned plants experiencing the greatest losses. However, both wake steering and axial induction are unable to mitigate the impact of upstream plants. Wake steering is constrained by plant geometry, since wake displacement is much less than the plant wake width, while axial induction requires curtailing the majority of turbines in upstream plants. Individual turbine strategies are limited by their effective scale and model representation. New wake models that include plant-scale physics are needed to facilitate the design of effective plant wake control strategies.

Scott, Ryan (ORCID:0000000328107574)↗
DOE OSTI2023

Large Eddy Simulation of Externally Induced Ingress about an Axial Seal by Stator Vanes

Turbine inlet temperatures in advanced gas turbines could be as high as 2000 °C. To prevent ingress of this hot gas into the wheelspace between the stator and rotor disks, whose metals can only handle temperatures up to 850 °C, rim seals and sealing flows are used. This study examines the abilities of large eddy simulation (LES) based on the WALE subgrid model and Reynolds-averaged Navier–Stokes (RANS) based on the SST model in predicting ingress in a rotor–stator configuration with vanes but no blades, a configuration with experimental data for validation. Results were obtained for an operating condition, where the ratio of the external Reynolds number to the rotational Reynolds number is 0.538. At this operating condition, both LES and RANS were found to correctly predict the coefficient of pressure, C p , located downstream of the vanes and upstream of the seal, but only LES was able to correctly predict the sealing effectiveness. This shows C p by itself is inadequate in quantifying externally induced ingress. RANS was unable to predict the sealing effectiveness because it significantly under predicted the pressure drop in the hot gas path along the axial direction, especially about the seal region. This affected the pressure difference across the seal in the radial direction, which ultimately drives ingress.

42 ENGINEERING↗
NASA NTRS1947

Analysis of Performance of Jet Engine from Characteristics of Components I : Aerodynamic and Matching Characteristics of Turbine Component Determined with Cold Air

The performance of the turbine component of an NACA research jet engine was investigated with cold air. The interaction and the matching of the turbine with the NACA eight-stage compressor were computed with the combination considered as a jet engine. The over-all performance of the engine was then determined. The internal aerodynamics were studied to the extent of investigating the performance of the first stator ring and its influence on the turbine performance. For this ring, the stream-filament method for computing velocity distribution permitted efficient sections to be designed, but the design condition of free-vortex flow with uniform axial velocities was not obtained.

Goldstein, Arthur W↗
NASA NTRS2013

Characterization of Swirl-Venturi Lean Direct Injection Designs for Aviation Gas-Turbine Combustion

Injector geometry, physical mixing, chemical processes, and engine cycle conditions together govern performance, operability and emission characteristics of aviation gas-turbine combustion systems. The present investigation explores swirl-venturi lean direct injection combustor fundamentals, characterizing the influence of key geometric injector parameters on reacting flow physics and emission production trends. In this computational study, a design space exploration was performed using a parameterized swirl-venturi lean direct injector model. From the parametric geometry, 20 three-element lean direct injection combustor sectors were produced and simulated using steady-state, Reynolds-averaged Navier-Stokes reacting computations. Species concentrations were solved directly using a reduced 18-step reaction mechanism for Jet-A. Turbulence closure was obtained using a nonlinear κ-ε model. Results demonstrate sensitivities of the geometric perturbations on axially averaged flow field responses. Output variables include axial velocity, turbulent kinetic energy, static temperature, fuel patternation and minor species mass fractions. Significant trends have been reduced to surrogate model approximations, intended to guide future injector design trade studies and advance aviation gas-turbine combustion research.

Combustion Chamber↗
DOE OSTI2023

Large Eddy Simulation of Rotationally Induced Ingress and Egress around an Axial Seal between Rotor and Stator Disks

In gas turbines, the hot gas exiting the combustor can have temperatures as high as 2000 °C, and some of this hot gas enter into the space between the stator and rotor disks (wheelspace). Since the entering hot gas could damage the disks, its ingestion must be minimized. This is carried out by rim seals and by introducing a cooler flow from the compressor (sealing flow) into the wheelspace. Ingress and egress into rim seals are driven by the stator vanes, the rotor and its rotation, and the rotor blades. This study focuses on the ingress and egress driven by the rotor and its rotation. This is carried out by performing wall-resolved large eddy simulation (LES) around an axial seal in a rotor–stator configuration without vanes and blades. Results obtained show the mechanisms by which the rotor and its rotation induce ingress, egress, and flow trajectories. Kelvin–Helmholtz instability was found to create a wavy shear layer and displacement thickness that produces alternating regions of high and low pressures around the rotor side of the seal. Vortex shedding on the backward-facing side of the seal and its impingement on the rotor side of the seal also produces alternating regions of high and low pressures. The locations of the alternating regions of high and low pressures were found to be statistically stationary and to cause ingress to start on the rotor side of the seal. Vortex shedding and recirculating flow in the seal clearance also cause ingress by entrainment. With the effects of the rotor and its rotation on ingress and egress isolated, this study enables the effects of stator vanes and rotor blades to be assessed.

42 ENGINEERING↗
NASA NTRS1981

Improved method for calculating transonic velocities on blade-to-blade stream surfaces of a turbomachine

A method was developed to improve the accuracy of an existing computer program used to calculate transonic velocities on a blade-to-blade surface of a turbomachine. The method eliminates problems encountered in obtaining solutions with the velocity gradient equation when large gradients in velocity occur through the blade row. With the improved method, results indicate that the transonic solution can be obtained by scaling the velocities obtained at the reduced mass flow rate where all velocities are subsonic thereby eliminating the need for a solution of the velocity gradient equation. Solutions obtained with the scaling method on a two dimensional compressor cascade and an axial turbine stator show good agreement with experimental data. The results obtained for the stationary blade rows and comparison of analytical results obtained with and without the present method suggest that the method will yield an improved solution for centrifugal compressor impellers.

Wood, J. R.↗
NASA NTRS1985

Bearing Tester Data Compilation Analysis, and Reporting and Bearing Math Modeling

The magnitude and direction of fluid induced torques and forces on the 57 mm bearing cage is considered to be a contributing factor in possible cage instabilities that can produce intermittent high heating in the bearing. Analyses of the fluid forces and torques are presented. Heat generated by viscous fluid work was estimated for two flow diverter configurations and a coolant flow of 10 lbs/sec to support the thermal evaluation of the LOX Bearing Materials Tester. Results of the analysis of the LOX turbopump turbine end bearings are discussed. Coolant velocities for the no. 4 LOX turbopump turbine end bearings were estimated as a function of shaft speed and coolant flow rate. Contact angles and track width data were developed for the 57 mm bearing as functions of shaft speed, and axial and radial loads. The Advanced Dynamics of Rolling Elements (ADORE) computer program was installed on the MSFC UNIVAC 1100 and a test case successfully run. Both the text output and the plotting output were verified. The Bearing Seal and Materials Tester - Test Condition Data Base was developed. The parametric analysis of the operating characteristics of the LOX turbopump pump end bearing using the 45 mm bearing thermal model was begun.

Source record↗
NASA NTRS1982

Laser anemometer measurements in an annular cascade of core turbine vanes and comparison with theory

Laser measurements were made in an annular cascade of stator vanes operating at an exit critical velocity ratio of 0.78. Velocity and flow angles in the blade to blade plane were obtained at every 10 percent of axial chord within the passage and at 1/2 axial chord downstream of the vanes for radial positions near the hub, mean and tip. Results are presented in both plot and tabulated form and are compared with calculations from an inviscid, quasi three dimensional computer program. The experimental measurements generally agreed well with these theoretical calculations, an indication of the usefulness of this analytic approach.

Goldman, L. J.↗