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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

A critical evaluation of a three-dimensional Navier-Stokes CFD as a tool to design supersonic turbine stages

Three-dimensional flow phenomena in a supersonic turbine blade row were studied numerically to evaluate CFD as a tool to design supersonic turbine stages. The details of the three-dimensional flow structure inside the supersonic turbine blade row and the overall aerodynamic performance at design and off-design conditions are analyzed and the results are compared between the experimental data and the numerical results.

Hah, C.↗
NASA NTRS2008

Fatigue Crack Growth Behavior Evaluation of Grainex Mar-M 247 for NASA's High Temperature, High Speed Turbine Seal Test Rig

The fatigue crack growth behavior of Grainex Mar-M 247 is evaluated for NASA s Turbine Seal Test Facility. The facility is used to test air-to-air seals primarily for use in advanced jet engine applications. Because of extreme seal test conditions of temperature, pressure, and surface speeds, surface cracks may develop over time in the disk bolt holes. An inspection interval is developed to preclude catastrophic disk failure by using experimental fatigue crack growth data. By combining current fatigue crack growth results with previous fatigue strain-life experimental work, an inspection interval is determined for the test disk. The fatigue crack growth life of the NASA disk bolt holes is found to be 367 cycles at a crack depth of 0.501 mm using a factor of 2 on life at maximum operating conditions. Combining this result with previous fatigue strain-life experimental work gives a total fatigue life of 1032 cycles at a crack depth of 0.501 mm. Eddy-current inspections are suggested starting at 665 cycles since eddy current detection thresholds are currently at 0.381 mm. Inspection intervals are recommended every 50 cycles when operated at maximum operating conditions.

Delgado, Irebert R.↗
NASA NTRS2006

Mach 4 Test Results of a Dual-Flowpath, Turbine Based Combined Cycle Inlet

An experimental study was conducted to evaluate the performance of a turbine based combined cycle (TBCC) inlet concept, consisting of a low speed turbojet inlet and high speed dual-mode scramjet inlet. The main objectives of the study were (1) to identify any interactions between the low and the high speed inlets during the mode transition phase in which both inlets are operating simultaneously and (2) to determine the effect of the low speed inlet operation on the performance of the high speed inlet. Tests were conducted at a nominal freestream Mach number of 4 using an 8 percent scale model representing a single module of a TBCC inlet. A flat plate was installed upstream of the model to produce a turbulent boundary layer which simulated the full-scale vehicle forebody boundary layer. A flowmeter/back pressure device, with remote actuation, was attached aft of the high speed inlet isolator to simulate the back pressure resulting from dual-mode scramjet combustion. Results indicate that the inlets did not interact with each other sufficiently to affect inlet operability. Flow spillage resulting from a high speed inlet unstart did not propagate far enough upstream to affect the low speed inlet. Also, a low speed inlet unstart did not cause the high speed inlet to unstart. The low speed inlet improved the performance of the high speed inlet at certain conditions by diverting a portion of the boundary layer generated on the forebody plate.

Albertson, Cindy w.↗
NASA NTRS2013

Advanced Shock Position Control for Mode Transition in a Turbine Based Combined Cycle Engine Inlet Model

A dual flow-path inlet system is being tested to evaluate methodologies for a Turbine Based Combined Cycle (TBCC) propulsion system to perform a controlled inlet mode transition. Prior to experimental testing, simulation models are used to test, debug, and validate potential control algorithms. One simulation package being used for testing is the High Mach Transient Engine Cycle Code simulation, known as HiTECC. This paper discusses the closed loop control system, which utilizes a shock location sensor to improve inlet performance and operability. Even though the shock location feedback has a coarse resolution, the feedback allows for a reduction in steady state error and, in some cases, better performance than with previous proposed pressure ratio based methods. This paper demonstrates the design and benefit with the implementation of a proportional-integral controller, an H-Infinity based controller, and a disturbance observer based controller.

Csank, Jeffrey T.↗
NASA NTRS1975

Performance of a catalytic reactor at simulated gas turbine combustor operating conditions

The performance of a catalytic reactor 12 cm in diameter and 17 cm long was evaluated at simulated gas turbine combustor operating conditions using premixed propane and air. Inlet temperatures of 600 and 800 K, pressures of 3 and 6 atm, and reference velocities of 9 to 30 m/s were tested. Data were taken for equivalence ratios as high as 0.43. The operating range was limited on the low-temperature side by very poor efficiency; the minimum exit temperature for good performance ranged from 1400 to 1600 K depending on inlet conditions. As exit temperatures were raised above this minimum, emissions of unburned hydrocarbons decreased, carbon monoxide emissions became generally less than 1 g CO/kg fuel, and nitrogen oxides were less than about 0.1 g NO2/kg fuel.

Anderson, D. N.↗
NASA NTRS1979

Thermal stress analysis of ceramic gas-path seal components for aircraft turbines

Stress and temperature distributions were evaluated numerically for a blade-tip seal system proposed for gas turbine applications. The seal consists of an abradable ceramic layer on metallic backing with intermediate layers between the ceramic layer and metal substrate. The most severe stresses in the seal, as far as failure is concerned, are tensile stresses at the top of the ceramic layer and shear and normal stresses at the layer interfaces. All these stresses reach their maximum values during the deceleration phase of a test engine cycle. A parametric study was carried out to evaluate the influence of various design parameters on these critical stress values. The influences of material properties and geometric parameters of the ceramic, intermediate, and backing layers were investigated. After the parametric study was completed, a seal system was designed which incorporated materials with beneficial elastic and thermal properties in each layer of the seal. An analysis of the proposed seal design shows an appreciable decrease in the magnitude of the maximum critical stresses over those obtained with earlier configurations.

Kennedy, F. E.↗
NASA NTRS1976

Experimental evaluation of two premixing-prevaporizing fuel injection concepts for a gas turbine catalytic combustor

A premixing-prevaporizing fuel system to be used with a catalytic combustor was evaluated for possible application in an automotive gas turbine. Spatial fuel distribution and degree of vaporization were measured using jet A fuel. Two types of air blast injectors were tested, a splash groove injector and a multiple jet cross stream injector. Air swirlers with vane angles of 15 deg and 30 deg were used to improve the spatial fuel distribution in a 12 cm diameter tubular rig. Distribution and vaporization measurements were made 35.5 cm downstream of the injector. The spatial fuel distribution was nearly uniform with the multiple jet contrastream injector and the splash-groove injector with a 30 deg air swirler. The vaporization was nearly 100 percent at an inlet air temperature of 600 K, and at 800 K inlet air temperature fuel oxidation reactions were observed. The total pressure loss was less than 0.5 percent of the total pressure for the multiple jet cross stream injector and the splash groove injector (without air swirler) and less than 1 percent for the splash groove with a 30 deg air swirler.

Tacina, R.↗
NASA NTRS1985

Design overview of fiber-reinforced superalloy composites for the Space Shuttle main engine

This preliminary design study evaluated the potential of fiber-reinforced superalloys (FRS) for hot-section components of Space Shuttle Main Engine turbopumps. Emphasis was placed on uncooled turbine blades, with a more limited evaluation of FRS turbine stator vanes. The study included FRS properties evaluation, current structural design capability, and preliminary design and structural analysis. In addition, key technology needs were identified, and a plan was generated to develop operational hardware for advanced versions of the SSME. Based on projections of design properties for FRS composites comprising 50 volume percent of W-4Re-0.38Hf-0.02C wire filaments in a ductile superalloy matrix, it was concluded that FRS turbine blades offer the potential of significant improved operating life and higher temperature capability over the MAR-M-246(Hf) (DS) blades currently used in the SSME.

Lewis, J. R.↗
NASA NTRS1985

Turbine Vane External Heat Transfer. Volume 1: Analytical and Experimental Evaluation of Surface Heat Transfer Distributions with Leading Edge Showerhead Film Cooling

Progress in predictive design capabilities for external heat transfer to turbine vanes was summarized. A two dimensional linear cascade (previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils) was used to examine the effect of leading edge shower head film cooling on downstream heat transfer. The data were used to develop and evaluate analytical models. Modifications to the two dimensional boundary layer model are described. The results were used to formulate and test an effective viscosity model capable of predicting heat transfer phenomena downstream of the leading edge film cooling array on both the suction and pressure surfaces, with and without mass injection.

Turner, E. R.↗
NASA NTRS1994

Design and aerodynamic performance evaluation of a high-work mixed flow turbine stage

As axial and radial turbine designs have been pushed to their aerothermodynamic and mechanical limits, the mixed-flow turbine (MFT) concept has been projected to offer performance and durability improvements, especially when ceramic materials are considered. The objective of this NASA/U.S. Army sponsored mixed-flow turbine (AMFT) program was to determine the level of performance attainable with MFT technology within the mechanical constraints of 1997 projected ceramic material properties. The MFT geometry is similar to a radial turbine, exhibiting a large radius change from inlet to exit, but differing in that the inlet flowpath is not purely radial, nor axial, but mixed; it is the inlet geometry that gives rise to the name 'mixed-flow'. The 'mixed' orientation of the turbine inlet offers several advantages over radial designs by allowing a nonzero inlet blade angle yet maintaining radial-element blades. The oblique inlet not only improves the particle-impact survivability of the design, but improves the aerodynamic performance by reducing the incidence at the blade inlet. The difficulty, however, of using mixed-flow geometry lies in the scarcity of detailed data and documented design experience. This paper reports the design of a MFT stage designed with the intent to maximize aerodynamic performance by optimizing design parameters such as stage reaction, rotor incidence, flowpath shape, blade shape, vane geometry, and airfoil counts using 2-D, 3-D inviscid, and 3-D viscous computational fluid dynamics code. The aerodynamic optimization was accomplished while maintaining mechanical integrity with respect to vibration and stress levels in the rotor. A full-scale cold-flow rig test was performed with metallic hardware fabricated to the specifications of the hot ceramic geometry to evaluate the stage performance.

Neri, Remo N.↗
NASA NTRS2012

A Turbine Based Combined Cycle Engine Inlet Model and Mode Transition Simulation Based on HiTECC Tool

An inlet system is being tested to evaluate methodologies for a turbine based combined cycle propulsion system to perform a controlled inlet mode transition. Prior to wind tunnel based hardware testing of controlled mode transitions, simulation models are used to test, debug, and validate potential control algorithms. One candidate simulation package for this purpose is the High Mach Transient Engine Cycle Code (HiTECC). The HiTECC simulation package models the inlet system, propulsion systems, thermal energy, geometry, nozzle, and fuel systems. This paper discusses the modification and redesign of the simulation package and control system to represent the NASA large-scale inlet model for Combined Cycle Engine mode transition studies, mounted in NASA Glenn s 10- by 10-Foot Supersonic Wind Tunnel. This model will be used for designing and testing candidate control algorithms before implementation.

Csank, Jeffrey T.↗
NASA NTRS2012

A Turbine Based Combined Cycle Engine Inlet Model and Mode Transition Simulation Based on HiTECC Tool

An inlet system is being tested to evaluate methodologies for a turbine based combined cycle propulsion system to perform a controlled inlet mode transition. Prior to wind tunnel based hardware testing of controlled mode transitions, simulation models are used to test, debug, and validate potential control algorithms. One candidate simulation package for this purpose is the High Mach Transient Engine Cycle Code (HiTECC). The HiTECC simulation package models the inlet system, propulsion systems, thermal energy, geometry, nozzle, and fuel systems. This paper discusses the modification and redesign of the simulation package and control system to represent the NASA large-scale inlet model for Combined Cycle Engine mode transition studies, mounted in NASA Glenn s 10-foot by 10-foot Supersonic Wind Tunnel. This model will be used for designing and testing candidate control algorithms before implementation.

Csank, Jeffrey↗