SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Lubricant, thrust bearing, and ring seal tests for supersonic turbine engine mainshaft lubrication system
The investigation was carried out to determine the optimum alloy within the Ni-16Cr-Al-Y2O3 system for use as a vane material in advanced aircraft gas turbine engines. Six alloys containing nominally 4%, 5% and 6% Al with Y2O3 levels of 0.8% and 1.2% were prepared by mechanical attrition. Six small-scale, rectangular extrusions were produced from each powder lot for property evaluation. The approximate temperatures for incipient melting were found to be 1658 K (2525 F), 1644 K (2500 F) and 1630 K (2475 F) for the 4%, 5% and 6% aluminum levels, respectively. With the exception of longitudinal crystallographic texture, the eight extrusions selected for extensive evaluation either exceeded or were close to mechanical property goals. Major differences between the alloys became apparent during dynamic oxidation testing, and in particular during the 1366 K (2000 F)/500 hour Mach 1 tests carried out by NASA-Lewis. An aluminum level of 4.75% was subsequently judged to be optimum based on considerations of dynamic oxidation resistance, susceptibility to thermal fatigue cracking and melting point.
A number of viscoplastic constitutive theories and a conventional constitutive theory are evaluated and compared in their ability to predict nonlinear stress-strain behavior in gas turbine engine components at elevated temperatures. Specific application of these theories is directed towards the structural analysis of combustor liners undergoing transient, cyclic, thermomechanical load histories. The combustor liner material considered in this study is Hastelloy X. The material constants for each of the theories (as a function of temperature) are obtained from existing, published experimental data. The viscoplastic theories and a conventional theory are incorporated into a general purpose, nonlinear, finite element computer program. Several numerical examples of combustor liner structural analysis using these theories are given to demonstrate their capabilities. Based on the numerical stress-strain results, the theories are evaluated and compared.
The results of studies concerning the high temperature corrosion resistance of ZrO2-Y2O3, ZrO2-MgO, and Ca2SiO4 plasma-sprayed coatings, which may be used as gas turbine engine thermal barriers, are reported. The coatings were evaluated in atmospheric burner rig and pressurized passage tests, using GT No. 2 fuel in pure form and with sodium, sulfur and vanadium corrosive impurities doping. It is found that, while the coatings performed well in both pressurized passage and burner rig tests with pure fuel chemical reactions between the ceramic coatings and combustion gases/condensates resulted in coating degradation with impure fuels. Chemical reactions between the ceramic coatings and vanadium compounds played a critical role in coating degradation.
When all the technology studies were done and the accompanying market analyses were complete, the conclusion was that it is indeed possible to reduce the cost of turbine engines by a factor of 3 using low-cost manufacturing techniques and increased production rates. In the interest of reducing engine cost, some performance was sacrificed. Yet we ended up with about a 20 percent predicted improvement in SFC over current technology turboprops. However, even this level of improvement does not match the low SFC of reciprocating powerplants--particularly those advanced concepts described earlier. The 20 percent better SFC and much lower weight of a turboprop does mean that if such a powerplant were installed in a resized small airplane, one could save between 10 and 30 percent fuel relative to existing recip engines, depending on different mission and airplane combinations. The price of the aircraft would go down about 15 percent in the case of a high powered single, or 25 percent in the case of a normal size twin. The operating costs would decrease about 10 percent in the case of the single, and as much as 35 percent in the case of the twin.
The effectiveness of propellant atomization with and without air injection in the combustion chamber nozzle of a gas turbine engine is studied. Test show that the startup and burning performance of these combustion chambers can be improved by using an injection during the mechanical propellant atomization process. It is shown that the operational range of combustion chambers can be extended to poorer propellant mixtures by combined air injection mechanical atomization of the propellant.
The suitability of wrought oxide dispersion strengthened (ODS) superalloy sheet for gas turbine engine combustor applications was evaluated. Incoloy MA 956 (FeCrAl base) and Haynes Developmental Alloy (HDA) 8077 (NiCrAl base) were evaluated. Preliminary tests showed both alloys to be potentially viable combustor materials, with neither alloy exhibiting a significant advantage over the other. Both alloys demonstrated a +167C (300 F) advantage of creep and oxidation resistance with no improvement in thermal fatigue capability compared to a current generation combustor alloy (Hastelloy X). MA956 alloy was selected for further demonstration because it exhibited better manufacturing reproducibility than HDA8077. Additional property tests were conducted on MA956. To accommodate the limited thermal fatigue capability of ODS alloys, two segmented, mechanically attached, low strain ODS combustor design concepts having predicted fatigue lives or = 10,000 engine cycles were identified. One of these was a relatively conventional louvered geometry, while the other involved a transpiration cooled configuration. A series of 10,000 cycle combustor rig tests on subscale MA956 and Hastelloy X combustor components showed no cracking, thereby confirming the beneficial effect of the segmented design on thermal fatigue capability. These tests also confirmed the superior oxidation and thermal distortion resistance of the ODS alloy. A hybrid PW2037 inner burner liner containing MA956 and Hastelloy X components was designed and constructed.
Light weight and potential high temperature capability of intermetallic compounds, such as the aluminides, and structural ceramics, such as the carbides and nitrides, make these materials attractive for gas turbine engine applications. In terms of specific fuel consumption and specific thrust, revolutionary improvements over current technology are being sought by realizing the potential of these materials through their use as matrices combined with high strength, high temperature fibers. The U.S. along with other countries throughout the world have major research and development programs underway to characterize these composites materials; improve their reliability; identify and develop new processing techniques, new matrix compositions, and new fiber compositions; and to predict their life and failure mechanisms under engine operating conditions. The status is summarized of NASA's Advanced High Temperature Engine Materials Technology Program (HITEMP) and the potential benefits are described to be gained in 21st century transport aircraft by utilizing intermetallic and ceramic matrix composite materials.
The suitability of wrought oxide dispersion strengthened (ODS) superalloy sheet for gas turbine engine combustor applications was evaluated. Two yttria (Y2O3) dispersion strengthened alloys were evaluated; Incoloy MA956 and Haynes Development Alloy (HDA) 8077 (NiCrAl base). Preliminary tests showed both alloys to be potentially viable combustor materials, with neither alloy exhibiting a significant advantage over the other. MA956 was selected as the final alloy based on manufacturing reproducibility for evaluation as a burner liner. A hybrid PW2037 inner burner liner containing MA956 and Hastelloy X components and using a louvered configuration was designed and constructed. The louvered configuration was chosen because of field experience and compatibility with the bill of material PW2037 design. The simulated flight cycle for the ground based engine tests consisted of 4.5 min idle, 1.5 min takeoff and intermediate conditions in a PW2037 engine with average uncorrected combustor exit temperature of 1527 C. Post test evaluation consisting of visual observations and fluorescent penetrant inspections was conducted after 500 cycles of testing. No loss of integrity in the burner liner was shown.
An analysis was conducted to assess methods for, and performance implications of, cooling the passages (tubes) of a pulse detonation-based combustor conceptually installed in the core of a gas turbine engine typical of regional aircraft. Temperature-limited material stress criteria were developed from common-sense engineering practice, and available material properties. Validated, one-dimensional, numerical simulations were then used to explore a variety of cooling methods and establish whether or not they met the established criteria. Simulation output data from successful schemes were averaged and used in a cycle-deck engine simulation in order to assess the impact of the cooling method on overall performance. Results were compared to both a baseline engine equipped with a constant-pressure combustor and to one equipped with an idealized detonative combustor. Major findings indicate that thermal loads in these devices are large, but potentially manageable. However, the impact on performance can be substantial. Nearly one half of the ideally possible specific fuel consumption (SFC) reduction is lost due to cooling of the tubes. Details of the analysis are described, limitations are presented, and implications are discussed.
Understanding the unsteady aspects of turbine rotor flowfields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with surface-mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in three respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, two independent unsteady data acquisition systems and fundamental signal processing approaches were used. Finally, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools will contribute to future turbine programs such as those for reusable launch vehicles.
Understanding the unsteady aspects of turbine rotor flow fields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with miniature surface mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in two respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. A three-dimensional unsteady Navier-Stokes analysis was also used to blindly predict the unsteady flow field in the turbine at the design operating conditions and at +15 degrees relative incidence to the first-stage rotor. The predicted time-averaged and unsteady pressure distributions show good agreement with the experimental data. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools are contributing significantly to current Space Launch Initiative turbine airflow test and blade surface pressure prediction efforts.
This publication is the second part of the three part report of the project entitled "A Fully Nonmetallic Gas Turbine Engine Enabled by Additive Manufacturing" funded by NASA Aeronautics Research Institute (NARI). The objective of this project was to conduct additive manufacturing to produce aircraft engine components by Fused Deposition Modeling (FDM), using commercially available polyetherimides-Ultem 9085 and experimental Ultem 1000 mixed with 10% chopped carbon fiber. A property comparison between FDM-printed and injection molded coupons for Ultem 9085, Ultem 1000 resin and the fiber-filled composite Ultem 1000 was carried out. Furthermore, an acoustic liner was printed from Ultem 9085 simulating conventional honeycomb structured liners and tested in a wind tunnel. Composite compressor inlet guide vanes were also printed using fiber-filled Ultem 1000 filaments and tested in a cascade rig. The fiber-filled Ultem 1000 filaments and composite vanes were characterized by scanning electron microscope (SEM) and acid digestion to determine the porosity of FDM-printed articles which ranged from 25 to 31%. Coupons of Ultem 9085, experimental Ultem 1000 composites and XH6050 resin were tested at room temperature and 400F to evaluate their corresponding mechanical properties. A preliminary modeling was also initiated to predict the mechanical properties of FDM-printed Ultem 9085 coupons in relation to varied raster angles and void contents, using the GRC-developed MAC/GMC program.
An analytical study was conducted to assess the impact of the use of broad specification fuels with reduced hydrogen content on the design, performance, durability, emissions and operational characteristics of combustors for commercial aircraft gas turbine engines. The study was directed at defining necessary design revisions to combustors designed for use of Jet A when such are operated on ERBS (Experimental Referee Broad Specification Fuel) which has a nominal hydrogen content of 12.8 percent as opposed to 13.7 percent in current Jet A. The results indicate that improvements in combustor liner cooling, and/or materials, and methods of fuel atomization will be required if the hydrogen content of aircraft gas turbine fuel is decreased.
Size exclusion chromatography (SEC) utilizing mu-Bondagel and mu-Styragel columns with a tetrahydrofuran mobile phase was used to determine the chemical degradation of lubricant samples from a gas-turbine engine test. A MIL-L-27502 candidate, ester-based lubricant was run in a J57-29 engine at a bulk oil temperature of 216 C. In general, the analyses indicated a progressive loss of primary ester, additive depletion, and formation of higher molecular weight material. An oil sample taken at the conclusion of the test showed a reversal of this trend because of large additions of new oil. The high-molecular-weight product from the degraded ester absorbed strongly in the ultraviolet region at 254 nanometers. This would indicate the presence of chromophoric groups. An analysis of a similar ester lubricant from a separate high-temperature bearing test yielded qualitatively similar results.
The following report represents a compendium of selected speaker presentation materials and observations made by Prof O. Pinkus at the NASA/ASME/Industry sponsored workshop entitled "Tribological Limitations in Gas Turbine Engines" held on September 15-17, 1999 in Albany, New York. The impetus for the workshop came from the ASME's Research Committee on Tribology whose goal is to explore new tribological research topics which may become future research opportunities. Since this subject is of current interest to other industrial and government entities the conference received cosponsorship as noted above. The conference was well attended by government, industrial and academic participants. Topics discussed included current tribological issues in gas turbines as well as the potential impact (drawbacks and advantages) of future tribological technologies especially foil air bearings and magnetic beatings. It is hoped that this workshop report may serve as a starting point for continued discussions and activities in oil-free turbomachinery systems.
The following report represents a compendium of selected speaker presentation materials and observations made by Prof. O. Pinkus at the NASA/ASME/Industry sponsored workshop entitled "Tribological Limitations in Gas Turbine Engines" held on September 15-17, 1999 in Albany, New York. The impetus for the workshop came from the ASME's Research Committee on tribology whose goal is to explore new tribological research topics which may become future research opportunities. Since this subject is of current interest to other industrial and government entities the conference received cosponsorship as noted above. The conference was well attended by government, industrial, and academic participants. Topics discussed included current tribological issues in gas turbines as well as the potential impact (drawbacks and advantages) of future tribological technologies especially foil air bearings and magnetic bearings. It is hoped that this workshop report may serve as a starting point for continued discussions and activities in oil-free turbomachinery systems.
This paper describes ongoing testing of an adaptive control method to suppress high frequency thermo-acoustic instabilities like those found in lean-burning, low emission combustors that are being developed for future aircraft gas turbine engines. The method called Adaptive Sliding Phasor Averaged Control, was previously tested in an experimental rig designed to simulate a combustor with an instability of about 530 Hz. Results published earlier, and briefly presented here, demonstrated that this method was effective in suppressing the instability. Because this test rig did not exhibit a well pronounced instability, a question remained regarding the effectiveness of the control methodology when applied to a more coherent instability. To answer this question, a modified combustor rig was assembled at the NASA Glenn Research Center in Cleveland, Ohio. The modified rig exhibited a more coherent, higher amplitude instability, but at a lower frequency of about 315 Hz. Test results show that this control method successfully reduced the instability pressure of the lower frequency test rig. In addition, due to a certain phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling, a dramatic suppression of the instability was achieved by focusing control on the second harmonic of the instability. These results and their implications are discussed, as well as a hypothesis describing the mechanism of intra-harmonic coupling.