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At least 109 records · Page 6

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NASA NTRS1950

Effect of fuel volatility on altitude starting limits of a turbojet engine

The effect of fuel volatility on altitude starting limits of an axial-flow-compressor-type turbojet engine was investigated using fuels with Reid vapor pressures of 1.1 and 5.4 pounds per square inch. At flight Mach numbers from 0.40 to 0.85, the AN-F-58 fuel allowed consistent windmilling at altitudes 2000 to 8000 feet higher than was obtained with the 1.1-pound Reid vapor pressure fuel. At a flight Mach number of 0.25, ignition could not be established at any altitude with the lower-volatility fuel.

COMBUSTION - IGNITION OF GASES↗

NASA NTRS1953

Altitude Operational Characteristics of Prototype J40-WE-8 Turbojet Engine

The altitude operational characteristics of the prototype J40-WE-8 turbojet engine were determined in the NACA Lewis altitude wind tunnel. The operational characteristics that were obtained include (1) compressor surge characteristics, (2) acceleration and deceleration rates, (3) steady-state windmilling-engine characteristics, (4) altitude ignition characteristics, and (5) the effect of two grades of fuel on steady-state engine performance at an altitude of 45,000 feet and a flight Mach number of 0.20. The altitude ignition characteristics determined include the effects of (1) ignition fuel flow, (2) two ignition procedures, (3) fuel temperatures, and (4) two grades of fuel on ignition characteristics. The compressor surge line, when presented as a function of compressor pressure ratio and corrected engine speed, was not affected by changes in flight condition and was independent of engine-inlet installation and of the manner in which surge was approached, rapidly or slowly. Also there was no effect of altitude or engine-inlet installation on the compressor surge recovery line when presented as a function of compressor pressure ratio and corrected engine speed.

FUELS - RELATION TO ENGINE PERFORMANCE↗

NASA NTRS1945

Determination of the Stability and Control Characteristics of a Tailless All-Wing Airplane Model with Sweepback in the Langley Free-Flight Tunnel

Force and flight tests were performance on an all-wing model with windmilling propellers. Tests were conducted with deflected and retracted flaps, with and without auxiliary vertical tail surfaces, and with different centers of gravity and trim coefficients. Results indicate serious reduction of stick-fixed longitudinal stability because of wing-tip stalling at high lift coefficient. Directional stability without vertical tail is undesirably low. Low effective dihedral should be maintained. Elevator and rudder control system is satisfactory.

WING FLAPS - DEFLECTION↗

NASA NTRS1932

Vertical descent of the autogiro

The purpose of this report is to show that only part of the system of rotating blades really is in the "windmill decelerating attitude" when the profile drag is sufficiently low. This particular part receives more torque from the air loads than can be absorbed by the profile drag. As a result thereof the rotating autogiro blade, when its torque is zero, is in part a propeller which functions in the "annular vortex attitude."

Bennett, J A J↗

NASA NTRS1993

Particle kinetic simulation of high altitude hypervelocity flight

In this grant period, the focus has been on enhancement and application of the direct simulation Monte Carlo (DSMC) particle method for computing hypersonic flows of re-entry vehicles. Enhancement efforts dealt with modeling gas-gas interactions for thermal non-equilibrium relaxation processes and gas-surface interactions for prediction of vehicle surface temperatures. Both are important for application to problems of engineering interest. The code was employed in a parametric study to improve future applications, and in simulations of aeropass maneuvers in support of the Magellan mission. Detailed comparisons between continuum models for internal energy relaxation and DSMC models reveals that several discrepancies exist. These include definitions of relaxation parameters and the methodologies for implementing them in DSMC codes. These issues were clarified and all differences were rectified in a paper (Appendix A) submitted to Physics of Fluids A, featuring several key figures in the DSMC community as co-authors and B. Haas as first author. This material will be presented at the Fluid Dynamics meeting of the American Physical Society on November 21, 1993. The aerodynamics of space vehicles in highly rarefied flows are very sensitive to the vehicle surface temperatures. Rather than require prescribed temperature estimates for spacecraft as is typically done in DSMC methods, a new technique was developed which couples the dynamic surface heat transfer characteristics into the DSMC flow simulation code to compute surface temperatures directly. This model, when applied to thin planar bodies such as solar panels, was described in AIAA Paper No. 93-2765 (Appendix B) and was presented at the Thermophysics Conference in July 1993. The paper has been submitted to the Journal of Thermophysics and Heat Transfer. Application of the DSMC method to problems of practical interest requires a trade off between solution accuracy and computational expense and limitations. A parametric study was performed and reported in AIAA Paper No. 93-2806 (Appendix C) which assessed the accuracy penalties associated with simulations of varying grid resolution and flow domain size. The paper was also presented at the Thermophysics Conference and will be submitted to the journal shortly. Finally, the DSMC code was employed to assess the pitch, yaw, and roll aerodynamics of the Magellan spacecraft during entry into the Venus atmosphere at off-design attitudes. This work was in support of the Magellan aerobraking maneuver of May 25-Aug. 3, 1993. Furthermore, analysis of the roll characteristics of the configuration with canted solar panels was performed in support of the proposed 'Windmill' experiment. Results were reported in AIAA Paper No. 93-3676 (Appendix D) presented at the Atmospheric Flight Mechanics Conference in August 1993, and were submitted to Journal of Spacecraft and Rockets.

Haas, Brian L.↗

NASA NTRS1996

Aspirating Seal Development: Analytical Modeling and Seal Test Rig

This effort is to develop large diameter (22 - 36 inch) Aspirating Seals for application in aircraft engines. Stein Seal Co. will be fabricating the 36-inch seal(s) for testing. GE's task is to establish a thorough understanding of the operation of Aspirating Seals through analytical modeling and full-scale testing. The two primary objectives of this project are to develop the analytical models of the aspirating seal system, to upgrade using GE's funds, GE's 50-inch seal test rig for testing the Aspirating Seal (back-to-back with a corresponding brush seal), test the aspirating seal(s) for seal closure, tracking and maneuver transients (tilt) at operating pressures and temperatures, and validate the analytical model. The objective of the analytical model development is to evaluate the transient and steady-state dynamic performance characteristics of the seal designed by Stein. The transient dynamic model uses a multi-body system approach: the Stator, Seal face and the rotor are treated as individual bodies with relative degrees of freedom. Initially, the thirty-six springs are represented as a single one trying to keep open the aspirating face. Stops (Contact elements) are provided between the stator and the seal (to compensate the preload in the fully-open position) and between the rotor face and Seal face (to detect rub). The secondary seal is considered as part of the stator. The film's load, damping and stiffness characteristics as functions of pressure and clearance are evaluated using a separate (NASA) code GFACE. Initially, a laminar flow theory is used. Special two-dimensional interpolation routines are written to establish exact film load and damping values at each integration time step. Additionally, other user-routines are written to read-in actual pressure, rpm, stator-growth and rotor growth data and, later, to transfer these as appropriate loads/motions in the system-dynamic model. The transient dynamic model evaluates the various motions, clearances and forces as the seals are subjected to different aircraft maneuvers: Windmilling restart; start-ground idle; ground idle-takeoff; takeoff-burst chop, etc. Results of this model show that the seal closes appropriately and does not ram into the rotor for all of the conditions analyzed. The rig upgrade design for testing Aspirating Seals has been completed. Long lead-time items (forgings, etc.) have been ordered.

Bharat Bagepalli↗

NASA NTRS2001

Simulation of Aircraft Engine Blade-Out Structural Dynamics

A primary concern of aircraft structure designers is the accurate simulation of the blade-out event and the subsequent windmilling of the engine. Reliable simulations of the blade-out event are required to insure structural integrity during flight as well as to guarantee successful blade-out certification testing. The system simulation includes the lost blade loadings and the interactions between the rotating turbomachinery and the remaining aircraft structural components. General-purpose finite element structural analysis codes such as MSC NASTRAN are typically used and special provisions are made to include transient effects from the blade loss and rotational effects resulting from the engine's turbomachinery. The present study provides the equations of motion for rotordynamic response including the effect of spooldown speed and rotor unbalance and examines the effects of these terms on a cantilevered rotor. The effect of spooldown speed is found to be greater with increasing spooldown rate. The parametric term resulting from the mass unbalance has a more significant effect on the rotordynamic response than does the spooldown term. The parametric term affects both the peak amplitudes as well as the resonant frequencies of the rotor.

Lawrence, Charles↗

NASA NTRS2001

Simulation of Aircraft Engine Blade-Out Structural Dynamics

Lawrence, Charles↗

NASA NTRS1954

An Investigation in the Ames 40- by 80-Foot Wind Tunnel of a YT-56A Turboprop Engine Incorporating a Decoupler and a Controlled-Feathering Device

An investigation of a decoupler and a controlled-feathering device incorporated with the YT-56A turboprop engine has been made to determine the effectiveness of these devices in reducing the high negative thrust (drag) which accompanies power failure of this type of engine. Power failures were simulated by fuel cut-off, both without either device free to operate, and with each device free to operate singly. The investigation was made through an airspeed range from 50 to 230 mph. It was found that with neither device free to operate, the drag levels realized after power failures at airspeeds above 170 mph would impose vertical tail loads higher than those allowable for the YC-130, the airplane for which the test power package was designed. These levels were reached in approximately one second. The maximum drag realized after power failure was not appreciably altered by the use of the decoupler although the decoupler did put a limit on the duration of the peak drag. The controlled-feathering device maintained a level of essentially zero drag after power failure. The use of the decoupler in the YT-56A engine complicates windmilling air-starting procedures and makes it necessary to place operating restrictions on the engine to assure safe flight at low-power conditions,

Rogallo, Vernon L.↗

NASA NTRS1947

Free-Spinning-Tunnel Tests of a 1/16-Scale Model of the Chance Vought XF5U-1 Airplane, TED No. NACA 2349

Spin tests of a 1/16-scale model of the Chance Vought XF5U-1 airplane have been performed in the Langley 20-foot free-spinning tunnel. The effect of control position and movement upon the erect and inverted spin and recovery characteristics ae well as the effects of propellers, of stability flaps, and of various revisions to the design configuration have been determined for the normal fighter loading. The investigation also included spin recovery parachute, tumbling, and pilot-escape tests. For the original design configuration, with or without windmilling propellers, the recovery characteristics of the model were considered unsatisfactory. Increasing the maximum upward deflection of the ailavators from 45 deg to 65 deg resulted in greatly improved recovery characteristics. Dimensional revisions to the original airplane configuration, which satisfactorily improved the general spin and recovery characteristics of the model, consisted of: (1) a supplementary vertical tail 34 inches by 59 inches (full-scale) attached to a boom 80 inches aft of the trailing edge of the airplane in the plane of symmetry, (2) a large semispan undersurface spoiler placed along the airplane quarter-chord line and opened on the outboard side in a spin, or (3) two additional vertical tails 64 inches by 52 inches (full-scale) located at the tips of the ailavators. A satisfactory parachute arrangement for emergency spin recovery from demonstration spins was found to be an arrangement consisting of a 13.3-foot parachute attached by a 30-foot towline to the arresting gear mast on the airplane and opened simultaneously with an 8-foot parachute on the outboard end of the wing attached by a 3-foot towline. Tests indicated that pilot escape from a spin would be extremely hazardous unless the pilot is mechanically ejected from the cockpit. Model tumbling tests indicated that the airplane would not tumble.

White, Richard P.↗

NASA NTRS1946

Effect of an Auxiliary Belly Fuel Tank on the Low-Speed Static Stability Characteristics of a 1/5-Scale Model of the Grumman XF8F-1 Airplane, TED No. NACA 2384

In.order to determine the aerodynamic effects of an auxiliary belly fuel tank on the Grumman F8F-1 airplane, a wind-tunnel investigation was made on a l/5 - scale model of the Grumman XF8F-1 airplane. Pitch and yaw tests were made with the model in the cruising and landing configurations for windmilling and take-off power conditions. Tuft studies and static-pressure measurements were also made to determine the flow characteristics in the region of the fuel tank. It was found that, at low speed, the auxiliary fuel tank test conditions, especially with power on in the landing configuration at high lift coefficients. The static directional stability was decreased for most test conditions, but the addition of a fairing between the fuselage and fuel tank improved the directional stability slightly in the power-on clean condition. The effective dihedral and lateral force were increased for most of the conditions tested. The tuft studies and pressure measurements indicated that the removal of the away braces would improve the.flow characteristics considerably in the region of the fuel tank end might also decrease the buffeting of the belly tank at high speeds.

Cook, Charles B.↗

NASA NTRS1947

Lateral Stability Characteristics of a 1/8.33-Scale Powered Model of the Republic XF-12 Airplane

The XF-12 airplane is a high-performance photo-reconnaissance aircraft designed for the Army Air Forces by the Republic Aviation Corporation. An investigation of a 1/8.33 - scale powered model was made in the Langley l9-foot pressure tunnel to obtain information relative to the aerodynamic design of the airplane. The model was tested with and without the original vertical tail. and with two revised tails. For the revised tail no. 1, the span of the original vertical .tail was increased about 15 percent and the portion of the vertical tail between the stabilizer and fuselage behind the rudder hinge line was allowed to deflect simultaneously with the main rudder. Revision no. 2 incorporated the increased span, but the lower rudder was locked in the neutral position. For all the tail arrangements investigated it was indicated that the airplane will possess positive effective dihedral and will be directionally stable regardless of flap or power condition. The rudder effectiveness is greater for the revised tails than for the original tail, but this is offset by the increase in directional stability caused by the revised tail. All the rudder arrangements appear inadequate in trimming out the resultant yawing moments at zero yaw in a take - off condition with the left-hand outboard propeller windmilling and the remaining engines developing take-off power.

Pepper, Edward↗

NASA NTRS1953

Investigation of the Spinning and Tumbling Characteristics of a 1/25-Scale Model of the Lockheed XFV-1 Airplane in the Langley 20-Foot Free-Spinning Tunnel

An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a l/23-scale model of the Lockheed XFV-1 airplane to determine the effects of control setting and movement upon the erect-spin and recovery characteristics for a range of airplane loading conditions. A windmilling propeller was simulated on the model for some of the tests. The investigation included determination of the size of tail parachute required for emergency recovery from demonstration spins. The tumbling tendencies of the model were also investigated. The results indicated that any erect or inverted spin obtained on the airplane will be satisfactorily terminated if recovery is attempted by full rudder reversal accompanied by simultaneous lateral and longitudinal movement of the stick to neutral, The model test results showed that an 11.5-foot flat-type tail parachute (drag coefficient approximately 0.73) with a 27.5-foot towline will be effective as an emergency spin-recovery device during demonstration spins of the airplane. The model results also indicate that the airplane will not tumble for any.loading condition indicated possible.

Lee, Henry A.↗

NASA NTRS2004

Solid oxidized fuel cells seals leakage setup and testing

As the world s reserves of fossil fuels are depleted, the U.S. Government, as well as other countries and private industries, is researching solutions for obtaining power, answers that would be more efficient and environmentally friendly. For a long time engineers have been trying to obtain the benefits of clean electric power without heavy batteries or pollution-producing engines. While some of the inventions proved to be effective (i.e. solar panels or windmills) their applications are limited due to dependency on the energy source (i.e. sun or wind). Currently, as energy concerns increase, research is being carried out on the development of a Solid Oxide Fuel Cell (SOFC). The United States government is taking a proactive role in expanding the technology through the Solid State Energy Conversion Alliance (SECA) Program, which is coordinated by the Department of Energy. into an electrical energy. This occurs by the means of natural tendency of oxygen and hydrogen to chemically react. While controlling the process, it is possible to harvest the energy given off by the reaction. SOFCs use currently available fossil fuels and convert a variety of those fuels with very high efficiency (about 40% more efficient than modem thermal power plants). At the same time they are almost entirely nonpolluting and due to their size they can be placed in remote areas. The main fields where the application of the fuel cells appears to be the most useful for are stationary energy sources, transportation, and military applications. structure and materials must be resolved. All the components must be operational in harsh environments including temperatures reaching 800 C and cyclic thermal- mechanical loading. Under these conditions, the main concern is the requirement for hermetic seals to: (1) prevent mixing of the fuel and oxidant within the stack, (2) prevent parasitic leakage of the fuel from the stack, (3) prevent contamination of the anode by air leaking into the stack, (4) electrically isolate the individual cells within the stack, and (5) mechanically bond the cell components. The sealing challenges are aggravated by the need to maintain hermetic boundaries between the different flow paths within the fuel cell throughout cycled operation. Within the timeframe of my tenure, the main objective is to assist in building a state-of-art test facility.

Bastrzyk, Marta B.↗

NASA NTRS2004

New Tool Released for Engine-Airframe Blade-Out Structural Simulations

Researchers at the NASA Glenn Research Center have enhanced a general-purpose finite element code, NASTRAN, for engine-airframe structural simulations during steady-state and transient operating conditions. For steady-state simulations, the code can predict critical operating speeds, natural modes of vibration, and forced response (e.g., cabin noise and component fatigue). The code can be used to perform static analysis to predict engine-airframe response and component stresses due to maneuver loads. For transient response, the simulation code can be used to predict response due to bladeoff events and subsequent engine shutdown and windmilling conditions. In addition, the code can be used as a pretest analysis tool to predict the results of the bladeout test required for FAA certification of new and derivative aircraft engines. Before the present analysis code was developed, all the major aircraft engine and airframe manufacturers in the United States and overseas were performing similar types of analyses to ensure the structural integrity of engine-airframe systems. Although there were many similarities among the analysis procedures, each manufacturer was developing and maintaining its own structural analysis capabilities independently. This situation led to high software development and maintenance costs, complications with manufacturers exchanging models and results, and limitations in predicting the structural response to the desired degree of accuracy. An industry-NASA team was formed to overcome these problems by developing a common analysis tool that would satisfy all the structural analysis needs of the industry and that would be available and supported by a commercial software vendor so that the team members would be relieved of maintenance and development responsibilities. Input from all the team members was used to ensure that everyone's requirements were satisfied and that the best technology was incorporated into the code. Furthermore, because the code would be distributed by a commercial software vendor, it would be more readily available to engine and airframe manufacturers, as well as to nonaircraft companies that did not previously have access to this capability.

Lawrence, Charles↗

NASA NTRS2007

TBCC Fan Stage Operability and Performance

NASA s Fundamental Aeronautics Program is investigating turbine-based propulsion systems for access to space because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. Studies performed under NASA s NGLT and the NASP High Speed Propulsion Assessment (HiSPA) program indicated a variable cycle turbofan/ramjet was the best configuration to satisfy access-to-space mission requirements because this configuration maximizes the engine thrust-to-weight ratio while minimizing frontal area. To this end, NASA and GE teamed to design a Mach 4 variable cycle turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10X), fan speed (7X), inlet mass flow (3.5X), inlet pressure (8X), and inlet temperature (3X). The primary goal of the fan stage was to provide a high pressure ratio level with good efficiency at takeoff through the mid range of engine operation, while avoiding stall and losses at the higher flight Mach numbers, without the use of variable inlet guide vanes. Overall fan performance and operability therefore requires major consideration, as competing goals at different operating points and aeromechanical issues become major drivers in the design. To mitigate risk of meeting the unique design requirements for the fan stage, NASA and GE teamed to design and build a 57% engine scaled fan stage to be tested in NASA s transonic compressor facility. The objectives of this test are to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off, 2) transition over large swings in fan bypass ratio, 3) transition from turbofan to ramjet, and 4) fan windmilling operation at high Mach flight conditions. In addition, the fan stage design was validated by performing pre-test CFD analysis using both GE proprietary and NASA s APNASA codes. Herein we will discuss 1) the fan stage design, 2) the experiment including the unique facility and instrumentation, and 3) the comparison of pre-test CFD analysis to initial aerodynamic test results for the baseline fan stage configuration. Measurements and pre-test analysis will be compared at 37%, 50%, 80%, 90%, and 100% of design speed to assess the ability of state-of-the-art design and analysis tools to meet the fan stage performance and operability requirements for turbine based propulsion for access to space.

Suder, Kenneth L.↗

NASA NTRS1954

Preliminary Altitude Performance Data for the J65-B3 Turbojet Engine at Reynolds Number Indices from 0.2 to 0.8

Altitude performance characteristics of the J65-B3 turbojet engine and its components were obtained at engine-inlet conditions corresponding to Reynolds number indices from 0.2 to 0.8 over a range of corrected engine speeds from 70 to 110 percent of rated speed. Engine operational limits up to an altitude of 75,000 feet together with ignition and windmilling characteristics were also obtained. The engine and component data are presented both in graphical and in tabulated form. The operational characteristics are presented in graphical form.

Braithwaite, W.M.↗

NASA NTRS1954

Preliminary Altitude Performance Data for the J71-A2 (X-26) Turbojet Engine

Data were obtained in an altitude test chamber for a range of altitudes from 20,000 to 58,000 feet at a flight Mach number of 0.9, and for several flight Mach numbers at an altitude of 45,000 feet. Data approximating sea-level operation are also included. Engine component performance data are presented in addition to windmilling, exhaust-nozzle, and ejector performance.

Useller, James W.↗