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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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Variability of flyover noise measures for repeated flights of turbojet and piston engine transport aircraft
Measurement of variability of aircraft noise during level flight flyovers
Effect of operating variables on pollutant emissions from aircraft turbine engine combustors.
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Impact resistance of fiber composite blades used in aircraft turbine engines
Resistance of advanced fiber reinforced epoxy matrix composite materials to ballistic impact was investigated as a function of impacting projectile characteristics, and composite material properties. Ballistic impact damage due to normal impacts, was classified as transverse (stress wave delamination and splitting), penetrative, or structural (gross failure). Steel projectiles were found to be gelatin ice projectiles in causing penetrative damage leading to reduced tensile strength. Gelatin and ice projectiles caused either transverse or structural damage, depending upon projectile mass and velocity. Improved composite transverse tensile strength, use of dispersed ply lay-ups, and inclusion of PRD-49-1 or S-glass fibers correlated with improved resistance of composite materials to transverse damage. In non-normal impacts against simulated blade shapes, the normal velocity component of the impact was used to correlate damage results with normal impact results. Stiffening the leading edge of simulated blade specimens led to reduced ballistic damage, while addition of a metallic leading edge provided nearly complete protection against 0.64 cm diameter steel, and 1.27 cm diameter ice and gelatin projectiles, and partial protection against 2.54 cm diameter projectiles of ice and gelatin.
Fretting in aircraft turbine engines
The problems created by fretting in turbine engines are discussed. The areas of greatest wear identified with the fan, compressor, and turbine blade mountings being the most critical items. Various methods for reducing or eliminating fretting in a turbine engine are described. Vacuum deposition of coatings by sputtering and ion plating are recommended as an economic method of applying thin films to inhibit fretting.
Data analysis study and performance evaluation of the scanning laser Doppler system
A simulation program which provided information on theoretically expected vortex spectra, evaluations of potential algorithms, and expected location accuracies for given scan patterns is presented. Field tests using an aircraft engine flow field and aircraft vortices during flyby tests were compared to the results of the simulation. From these studies, a vortex location algorithm was developed which provided vortex location for one or two vortices as a function of time. Results of this algorithm used on data from flyby tests were used to study vortex transport, to evaluate system performance, and to provide suggestions for real-time vortex location algorithms. The results of real-time analysis were compared to those which were expected based on theoretical considerations.
Fretting in aircraft turbine engines
Fretting usually produces both polished and pitted areas. Fretting occurs in connection with the relative motion between contacting surface elements. Particular attention is given to fretting in seals, splines, fans, compressors, and turbine blades. Recommended approaches to mitigate fretting problems include a reduction of the adhesion between two surfaces. Approaches for doing this may include the development of alloys which spontaneously generate a self-healing surface film, different in composition from the bulk alloy. Other fundamental adhesion studies have shown that surface crystal structure and crystallographic orientation significantly affect the adhesion properties of alloys.
Jet-exit and airframe interference studies on twin-engine-fuselage aircraft installations
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Improved turbine disk design to increase reliability of aircraft jet engines
An analytical study was conducted on a bore entry cooled turbine disk for the first stage of the JT8D-17 high pressure turbine which had the potential to improve disk life over existing design. The disk analysis included the consideration of transient and steady state temperature, blade loading, creep, low cycle fatigue, fracture mechanics and manufacturing flaws. The improvement in life of the bore entry cooled turbine disk was determined by comparing it with the existing disk made of both conventional and advanced (Astroloy) disk materials. The improvement in crack initiation life of the Astroloy bore entry cooled disk is 87% and 67% over the existing disk made of Waspaloy and Astroloy, respectively. Improvement in crack propagation life is 124% over the Waspaloy and 465% over the Astroloy disks. The available kinetic energies of disk fragments calculated for the three disks indicate a lower fragment energy level for the bore entry cooled turbine disk.
An improved turbine disk design to increase reliability of aircraft jet engines
An analytical study was performed on a novel disk design to replace the existing high-pressure turbine, stage 1 disk on the CF6-50 turbofan engine. Preliminary studies were conducted on seven candidate disk design concepts. An integral multidisk design with bore entry of the turbine blade cooling air was selected as the improved disk design. This disk has the unique feature of being redundant such that if one portion of the disk would fail, the remaining portion would prevent the release of large disk fragments from the turbine system. Low cycle fatigue lives, initial defect propagation lives, burst speed, and the kinetic energies of probable disk fragment configurations were calculated, and comparisons were made with the existing disk, both in its current material, IN 718, and with the substitution of an advanced alloy, Rene 95. The design for redundancy approach which necessitated the addition of approximately 44.5 kg (98 lb) to the design disk substantially improved the life of the disk. The life to crack initiation was increased from 30,000 cycles to more than 100,000 cycles. The cycles to failure from initial defect propagation were increased from 380 cycles to 1564 cycles. Burst speed was increased from 126 percent overspeed to 149 percent overspeed. Additionally, the maximum fragment energies associated with a failure were decreased by an order of magnitude.
Trends in high temperature materials technology for advanced aircraft turbine engines
The utilization of high temperature materials for turbine blades, disks, vanes, and combustors is reviewed. The use of directionally solidified eutectic alloys and of metal fiber reinforced superalloys as turbine blade materials is discussed. The application of powder metallurgy to the production of turbine disks is considered. It is shown that oxide-dispersion strengthened alloys (including NiCr and NiCr-Al) and ceramics are the best refractory turbine vane materials. The use of heat resistant sheet alloys as combustor materials is examined. The cost benefits to be derived from the application of high temperature technology to advanced CTOL and STOL engines are described.
Cost/benefit analysis of advanced materials technologies for future aircraft turbine engines
The cost/benefits of advance commercial gas turbine materials are described. Development costs, estimated payoffs and probabilities of success are discussed. The materials technologies investigated are: (1) single crystal turbine blades, (2) high strength hot isostatic pressed turbine disk, (3) advanced oxide dispersion strengthened burner liner, (4) bore entry cooled hot isostatic pressed turbine disk, (5) turbine blade tip - outer airseal system, and (6) advance turbine blade alloys.
TCM aircraft piston engine emission reduction program
The technology necessary to safely reduce general aviation piston engine exhaust emissions to meet the EPA 1980 Emission Standards with minimum adverse effects on cost, weight, fuel economy, and performance was demonstrated. A screening and assessment of promising emission reduction concepts was provided, and the preliminary design and development of those concepts was established. A system analysis study and a decision making procedure were used by TCM to evaluate, trade off, and rank the candidate concepts from a list of 14 alternatives. Cost, emissions, and 13 other design criteria considerations were defined and traded off against each candidate concept to establish its merit and emission reduction usefulness. A computer program was used to aid the evaluators in making the final choice of three concepts.
Minimum-time acceleration of aircraft turbofan engines
Minimum-time accelerations of the F100 turbofan engine are presented. A piecewise-linear engine model, having three state variables and four control variables, is used to obtain the minimum-time solutions. The linear model which applies at a given time in the trajectory is determined by calculating a normalized distance from the current state to the equilibrium state associated with each linear model. The linear model associated with the closest equilibrium point is then used. The control histories for the minimum-time solutions are used as input to a nonlinear simulation of the F100 engine to verify the accuracy of the piecewise linear solutions.
Dual cycle aircraft turbine engine
A method and apparatus are presented for improving operating efficiency over broad ranges of flight conditions and for reducing jet engine noise output in takeoff and landing by controlling the airflow entering and exiting the engines. A turbojet engine apparatus is described which operates efficiently at both subsonic and supersonic speeds and a method is described which enables a turbofan with an associated satellite turbojet or turbofan to operate more efficiently at both subsonic and supersonic speeds. In both cases, takeoff and landing noise is reduced substantially. The apparatus consists essentially of arranging for two separate portions of an engine to act upon one airstream or, alternately, to operate on independent airstreams.
Comparison of Estimated with Measured Maximum Instantaneous Distortion Using Flight Data from an Axisymmetric Mixed Compression Inlet
YF-12C flight-measured inlet dynamic distortion data are compared with predictions made on the basis of the method reported by Melick et al. (1976). The YF-12C aircraft is a twin engine aircraft capable of speeds above 3. The inlets have a translating spike to control the inlet throat area. A bypass system is used to control the terminal shock of the inlet for operation in the mixed compression mode. The dynamic data were obtained with the aid of 24 high frequency response total pressure sensors. The model of Melick et al. is discussed along with the computer program used to implement the model. It is found that the predictions of maximum instantaneous distortion are within 20 percent of the measured values, which had been obtained at Mach numbers of 1.8, 2.1, 2.5, and 3.0.
Minimum-time acceleration of aircraft turbofan engines
Minimum-time accelerations of the F100 turbofan engine are presented. A piecewise-linear engine model, having three state variables and four control variables, is used to obtain the minimum-time solutions. The linear model which applies at a given time in the trajectory is determined by calculating a normalized 'distance' from the current state to the equilibrium state associated with each linear model. The linear model associated with the closest equilibrium point is then used. The control histories for the minimum-time solutions are used as input to a nonlinear simulation of the F100 engine to verify the accuracy of the piecewise-linear solutions.
Structural analysis of hollow blades: Torsional stress analysis of hollow fan blades for aircraft jet engines
A torsional stress analysis of hollow fans blades by the finite element method is presented. The fans are considered to be double circular arc blades, hollowed 30 percent, and twisted by a component of the centrifugal force by the rated revolution. The effects of blade hollowing on strength and rigidity are discussed. The effects of reinforcing webs, placed in the hollowed section in varying numbers and locations, on torsional rigidity and the convergence of stresses, are reported. A forecast of the 30 percent hollowing against torsional loadings is discussed.