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At least 37 records · Page 2

All sources NASA NTRS DOE OSTI

NASA NTRS1982

Cold-air performance of a 15.41-cm-tip-diameter axial-flow power turbine with variable-area stator designed for a 75-kW automotive gas turbine engine

An experimental evaluation of the aerodynamic performance of the axial flow, variable area stator power turbine stage for the Department of Energy upgraded automotive gas turbine engine was conducted in cold air. The interstage transition duct, the variable area stator, the rotor, and the exit diffuser were included in the evaluation of the turbine stage. The measured total blading efficiency was 0.096 less than the design value of 0.85. Large radial gradients in flow conditions were found at the exit of the interstage duct that adversely affected power turbine performance. Although power turbine efficiency was less than design, the turbine operating line corresponding to the steady state road load power curve was within 0.02 of the maximum available stage efficiency at any given speed.

Mclallin, K. L.↗

NASA NTRS1982

Cost/benefit studies of advanced materials technologies for future aircraft turbine engines: Materials for advanced turbine engines

Cost benefit studies were conducted on six advanced materials and processes technologies applicable to commercial engines planned for production in the 1985 to 1990 time frame. These technologies consisted of thermal barrier coatings for combustor and high pressure turbine airfoils, directionally solidified eutectic high pressure turbine blades, (both cast and fabricated), and mixers, tail cones, and piping made of titanium-aluminum alloys. A fabricated titanium fan blisk, an advanced turbine disk alloy with improved low cycle fatigue life, and a long-life high pressure turbine blade abrasive tip and ceramic shroud system were also analyzed. Technologies showing considerable promise as to benefits, low development costs, and high probability of success were thermal barrier coating, directionally solidified eutectic turbine blades, and abrasive-tip blades/ceramic-shroud turbine systems.

Stearns, M.↗

NASA NTRS1982

Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 2: Stage performance

The aerodynamic performance of the compressor-drive turbine of the DOE upgraded gas turbine engine was determined in low temperature air. The as-received cast rotor blading had a significantly thicker profile than design and a fairly rough surface finish. Because of these blading imperfections a series of stage tests with modified rotors were made. These included the as-cast rotor, a reduced-roughness rotor, and a rotor with blades thinned to near design. Significant performance changes were measured. Tests were also made to determine the effect of Reynolds number on the turbine performance. Comparisons are made between this turbine and the compressor-drive turbine of the DOE baseline gas turbine engine.

Roelke, R. J.↗

NASA NTRS1957

Component Performance Investigation of J71 Experimental Turbine VII: Effect of First-stator Adjustment; Over-all Performance of J71-97 Turbine with 87-percent-design Stator Area

The performance results, as well as a compressor-turbine match point study based on a particular mode of engine operation, are compared with those previously obtained with the same turbine having first-stator areas 70, 95.6, and 132 percent of design. Maximum efficiencies for all turbine configurations were high, on the order of 87 to 89 percent. The efficiencies obtained at the match points for the 87-, 95.6-, and 132-percent turbines were slightly lower. No match point for the 70-percent turbine was obtained because, with first-stator areas less than 75.5 percent of design, the turbine would develop insufficient work to drive the compressor at the match point.

Schum, Harold J↗

NASA NTRS1947

Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine: Compressor and Turbine Performance Characteristics - 4

As part of an investigation of the performance and operational characteristics of the TG-100A gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100R. The highest compressor pressure ratio was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Wallner, Lewis E.↗

NASA NTRS1947

Cooling of Gas Turbines I - Effects of Addition of Fins to Blade Tips and Rotor, Admission of Cooling Air Through Part of Nozzles, and Change in Thermal Conductivity of Turbine Components

An analysis was developed for calculating the radial temperature distribution in a gas turbine with only the temperatures of the gas and the cooling air and the surface heat-transfer coefficient known. This analysis was applied to determine the temperatures of a complete wheel of a conventional single-stage impulse exhaust-gas turbine. The temperatures were first calculated for the case of the turbine operating at design conditions of speed, gas flow, etc. and with only the customary cooling arising from exposure of the outer blade flange and one face of the rotor to the air. Calculations were next made for the case of fins applied to the outer blade flange and the rotor. Finally the effects of using part of the nozzles (from 0 to 40 percent) for supplying cooling air and the effects of varying the metal thermal conductivity from 12 to 260 Btu per hour per foot per degree Farenheit on the wheel temperatures were determined. The gas temperatures at the nozzle box used in the calculations ranged from 1600F to 2000F. The results showed that if more than a few hundred degrees of cooling of turbine blades are required other means than indirect cooling with fins on the rotor and outer blade flange would be necessary. The amount of cooling indicated for the type of finning used could produce some improvement in efficiency and a large increase in durability of the wheel. The results also showed that if a large difference is to exist between the effective temperature of the exhaust gas and that of the blade material, as must be the case with present turbine materials and the high exhaust-gas temperatures desired (2000F and above), two alternatives are suggested: (a) If metal with a thermal conductivity comparable with copper is used, then the blade temperature can be reduced by strong cooling at both the blade tip and root. The center of the blade will be less than 2000F hotter than the ends; (b) With low conductivity materials some method of direct cooling other than partial admission of cooling air is essential. From this study, it can be deduced that indirect cooling of turbine blades will not make possible large increases in gas temperature.

Brown, Byron↗

NASA NTRS1948

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine: Compressor and Turbine Performance Characteristics - 4

As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Wallner, Lewis E.↗

NASA NTRS2011

Capability Extension to the Turbine Off-Design Computer Program AXOD With Applications to the Highly Loaded Fan-Drive Turbines

The axial flow turbine off-design computer program AXOD has been upgraded to include the outlet guide vane (OGV) into its acceptable turbine configurations. The mathematical bases and the techniques used for the code implementation are described and discussed in lengths in this paper. This extended capability is verified and validated with two cases of highly loaded fan-drive turbines, designed and tested in the V/STOL Program of NASA. The first case is a 4 1/2-stage turbine with an average stage loading factor of 4.66, designed by Pratt & Whitney Aircraft. The second case is a 3 1/2-stage turbine with an average loading factor of 4.0, designed in-house by the NASA Lewis Research Center (now the NASA Glenn Research Center). Both cases were experimentally tested in the turbine facility located at the Glenn Research Center. The processes conducted in these studies are described in detail in this paper, and the results in comparison with the experimental data are presented and discussed. The comparisons between the AXOD results and the experimental data are in excellent agreement.

Chen, Shu-cheng S.↗

NASA NTRS2016

Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

Gas turbine compressor and turbine blade tip clearance (i.e., the radial distance between the blade tip of an axial compressor or turbine and the containment structure) is a major contributing factor to gas path sealing, and can significantly affect engine efficiency and operational temperature. This paper details the creation of a generic but realistic high pressure turbine tip clearance model that may be used to facilitate active tip clearance control system research. This model uses a first principles approach to approximate thermal and mechanical deformations of the turbine system, taking into account the rotor, shroud, and blade tip components. Validation of the tip clearance model shows that the results are realistic and reflect values found in literature. In addition, this model has been integrated with a gas turbine engine simulation, creating a platform to explore engine performance as tip clearance is adjusted. Results from the integrated model explore the effects of tip clearance on engine operation and highlight advantages of tip clearance management.

aerothermodynamics↗

NASA NTRS2016

Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

gas path dynamics↗

NASA NTRS1952

Investigation of turbines suitable for use in a turbojet engine with high compressor pressure ratio and low compressor-tip speed I : turbine-design requirements for several engine operating conditions

In order to insure satisfactory turbine performance under five turbojet-engine operating conditions, turbine design requirements were determined for the following operating conditions: take-off, maximum thrust at altitude, altitude cruising at rated rotative speed, altitude cruising with maximum-thrust exhaust nozzle area, and engine acceleration at 80 percent equivalent design rotative speed. If cruising is to be at rated engine speed, both cruising and take-off should be considered in turbine design. Design requirements for the other conditions are nearly identical. Without compressor-exit bleed, a turbine cannot accelerate this compressor at 80 percent of rated speed; with 28.6-percent bleed, the design requirements are within those for take-off.

TURBINES--AXIAL FLOW↗

NASA NTRSDate not supplied

Numerical Investigation of Combustor-Turbine Interactions with a Two-Stage High-Pressure Turbine Including Cooling Airflows

Combustor–turbine interactions (CTIs) are investigated by performing three-dimensional unsteady simulations using a realistic combustor and high-pressure turbine (HPT) configurations from the Energy Efficient Engine (E3) program. To understand CTIs, we compare the numerically predicted flow fields from single-component simulations (Step 1: the combustor + the first-stage stator of turbine; Step 2: the two-stage HPT imposing the time-averaged flow solution from Step 1 as the inflow boundary condition) and a fully coupled combustor–turbine simulation (Step 3) at the sea-level takeoff (SLTO) condition. In addition to three previous simulations where the cooling airflows inside the HPT had been neglected for all Step 1, Step 2, and Step 3, two new simulations of Step 2 and Step 3 take into account the cooling airflows using the source team approach. In this approach, to mimic the cooling airflows, we impose the source term at a specific area of a cooling airflow hole at the solid surfaces without making a mesh of each hole. The objective of this study is twofold. One is to investigate the effect of the presence of the cooling airflows on the aerodynamics of the combustor and HPT as well as the HPT performance. The second is to perform a detailed comparison among the calculated flowfields by two different numerical schemes, the central-difference with the standard Jameson–Schmidt–Turkel (CD-JST) scheme and the AUSM scheme. There is a noticeable difference in the hot-streak distributions at the first-stage stator. In addition, depending on the choice of the numerical scheme and the presence of the cooling airflows, an occurrence of shock waves appearing at the first-stage stator is greatly influenced. Thus, this has a noticeable impact on the HPT performance. It is shown that the estimated turbine efficiencies from Step 3 are about 7 % less than the ones from Step 2.

Combustor-turbine interaction↗

NASA NTRSDate not supplied

Numerical Investigation of Combustor-Turbine Interactions with Cooling Air Flows Included in High-Pressure Turbine

In this paper, we would like to include the cooling airflows inside the turbine and perform the fully-coupled combustor–turbine simulation at a realistic sea-level takeoff (SLTO) condition. The new results are compared with the previous study,in which we performed the fully-coupled combustor–turbine simulation without the cooling airflows and the purge flows inside the HPT. This paper also serves to provide a better understanding about the effect of the unsteadiness in the turbine inflow condition on the turbine performance in the presence of the cooling airflows.

LES↗

NASA NTRS1974

Experimental investigation of a 4 and 1/2 stage turbine with very high stage loading factor. 2: Turbine performance

The experimental test program results of a 4 1/2-stage turbine with a very high stage loading factor are presented. A four-stage turbine was tested with and without outlet turning vanes. The 4 1/2-stage turbine achieved a design point total-to-total efficiency of 0.853. The outlet turning vane design point performance was 0.4 percent of the overall 4 1/2-stage turbine efficiency. Tests were conducted at various levels of Reynolds number and indicated decreases in turbine efficiency and equivalent weight flow with decreasing Reynolds number.

Walker, N. D.↗

NASA NTRS1974

Cold-air investigation of a turbine for high temperature-engine application. 5: Two-stage turbine performance as affected by variable stator area

The stator areas of the design two-stage turbine were both decreased and increased by nominally 30 percent, and the performances of the two turbines are compared with that of the design stator area turbine. Turbine efficiency decreased with stator area changes. Closing the stator area resulted in the more severe efficiency loss. The decrease in efficiency for both turbines is attributable to rotor incidence, off-design blade-surface velocities, and adverse reaction changes across the blade rows.

Behning, F. P.↗

NASA NTRS1984

Extended parametric representation of compressor fans and turbines. Volume 2: Part user's manual (parametric turbine)

A turbine modeling technique has been developed which will enable the user to obtain consistent and rapid off-design performance from design point input. This technique is applicable to both axial and radial flow turbine with flow sizes ranging from about one pound per second to several hundred pounds per second. The axial flow turbines may or may not include variable geometry in the first stage nozzle. A user-specified option will also permit the calculation of design point cooling flow levels and corresponding changes in efficiency for the axial flow turbines. The modeling technique has been incorporated into a time-sharing program in order to facilitate its use. Because this report contains a description of the input output data, values of typical inputs, and example cases, it is suitable as a user's manual. This report is the second of a three volume set. The titles of the three volumes are as follows: (1) Volume 1 CMGEN USER's Manual (Parametric Compressor Generator); (2) Volume 2 PART USER's Manual (Parametric Turbine); (3) Volume 3 MODFAN USER's Manual (Parametric Modulation Flow Fan).

Coverse, G. L.↗

NASA NTRS2003

Calculation of Turbine Axial Thrust by Coupled CFD Simulations of the Main Flow Path and Secondary Cavity Flow in an SLI LOX Turbine

Each single reusable Space Launch Initiative (SLI) booster rocket is an engine operating at a record vacuum thrust level of over 730,000 Ibf using LOX and LH2. This thrust is more than 10% greater than that of the Delta IV rocket, resulting in relatively large LOX and LH2 turbopumps. Since the SLI rocket employs a staged combustion cycle the level of pressure is very high (thousands of psia). This high pressure creates many engineering challenges, including the balancing of axial-forces on the turbopumps. One of the main parameters in the calculation of the axial force is the cavity pressure upstream of the turbine disk. The flow in this cavity is very complex. The lack of understanding of this flow environment hinders the accurate prediction of axial thrust. In order to narrow down the uncertainty band around the actual turbine axial force, a coupled, unsteady computational methodology has been developed to simulate the interaction between the turbine main flow path and the cavity flow. The CORSAIR solver, an unsteady three- dimensional Navier-Stokes code for turbomachinery applications, was used to solve for both the main and the secondary flow fields. Turbine axial thrust values are presented in conjunction with the CFD simulation, together with several considerations regarding the turbine instrumentation for axial thrust estimations during test.

Dorney, D. J.↗

NASA NTRS1950

Analytical investigation of turbines with adjustable stator blades and effect of these turbines on jet-engine performance

Adjustable-stator turbines are applied to turbojet engines and probable performance is compared with conventional engines with and without variable-area exhaust nozzles. Variation in stator-exit angle and exhaust area was not excessive for wide ranges of engine output. Variation in turbine efficiency for contemporary turbines equipped with adjustable stators was small. Improvements from 4.5 to 17 percent in over-all engine specific fuel consumption over conventional engines and from 2 to 8.5 percent over engines equipped with only adjustable-area exhaust nozzles were obtained at 60-percent rated power with adjustable-stator turbines and variable-area exhaust nozzles. The improvements depend on design parameters.

ENGINE TYPES, COMPARISON↗