Interrelations between combustion phenomena and mechanical properties in solid propellant rocket motors.
Static and dynamic combustion phenomena effects on grain structural design for solid propellant rocket engine
PROPELLANT GRAIN↗
SEARCH · Search NASA
Results for “static engine”
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.
SUMS calibration test report
Calibration was performed on the shuttle upper atmosphere mass spectrometer (SUMS). The results of the calibration and the as run test procedures are presented. The output data is described, and engineering data conversion factors, tables and curves, and calibration on instrument gauges are included. Static calibration results which include: instrument sensitive versus external pressure for N2 and O2, data from each scan of calibration, data plots from N2 and O2, and sensitivity of SUMS at inlet for N2 and O2, and ratios of 14/28 for nitrogen and 16/32 for oxygen are given.
Robertson, G.↗
The 1991 International Aerospace and Ground Conference on Lightning and Static Electricity, volume 2
The proceedings of the conference are reported. The conference focussed on lightning protection, detection, and forecasting. The conference was divided into 26 sessions based on research in lightning, static electricity, modeling, and mapping. These sessions spanned the spectrum from basic science to engineering, concentrating on lightning prediction and detection and on safety for ground facilities, aircraft, and aerospace vehicles.
Source record↗
Structural dynamic and aeroelastic considerations for hypersonic vehicles
Structural dynamic and aeroelastic considerations applicable to hypersonic vehicles are discussed. Emphasis is given to aerospace plane configurations. The definition of aerothermoelasticity and the operational flight environment are reviewed, and structural dynamic and aeroelastic areas of concern are individually discussed, including vibration, landing and taxiing, propellant dynamics, acoustics, lifting surface flutter, panel flutter, control surface buzz, buffeting, gust response, and static aeroelasticity. Recent research results from all-moveable delta-wing aerolastic studies, engine inlet lip aeroelastic analysis, and studies of thermal effects on vibration frequencies, aerodynamic heating effects on flutter, and active control of aeroelastic response are reviewed.
Cazier, F. W., Jr.↗
Flight Behavior of the X-2 Research Airplane to a Mach Number of 3.20 and a Geometric Altitude of 126,200 Feet
The maximum Mach number and altitude capabilities of the Bell X-2 research airplane were achieved during a program conducted by the U.S. Air Force with Bell Aircraft Corp. providing operational support and the National Aeronautics and Space Administration providing instrumentation and advisory engineering assistance. A maximum geometric altitude of 126,200 feet was attained at a static pressure of 9.4 pounds per square foot and a dynamic pressure of 19.1 pounds per square foot. During the last flight of the airplane, a maximum Mach number of 3.20 was reached. The directionally divergent maneuver which terminated the final high Mach number flight was precipitated by the loss in directional stability that resulted from increasing the angle of attack. The yawing moment from the lateral control was sufficient to initiate the divergence and also to cause,, indirectly, rolling moments that were greater than the aileron capabilities of the airplane. The ensuing violent motions-resulting from inertial roll coupling caused the loss of the aircraft.
Day, Richard E.↗
Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program
The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems.
Asbury, Scott C.↗
On the Exit Boundary Condition for One-Dimensional Calculations of Pulsed Detonation Engine Performance
In one-dimensional calculations of pulsed detonation engine (PDE) performance, the exit boundary condition is frequently taken to be a constant static pressure. In reality, for an isolated detonation tube, after the detonation wave arrives at the exit plane, there will be a region of high pressure, which will gradually return to ambient pressure as an almost spherical shock wave expands away from the exit, and weakens. Initially, the flow is supersonic, unaffected by external pressure, but later becomes subsonic. Previous authors have accounted for this situation either by assuming the subsonic pressure decay to be a relaxation phenomenon, or by running a two-dimensional calculation first, including a domain external to the detonation tube, and using the resulting exit pressure temporal distribution as the boundary condition for one-dimensional calculations. These calculations show that the increased pressure does affect the PDE performance. In the present work, a simple model of the exit process is used to estimate the pressure decay time. The planar shock wave emerging from the tube is assumed to transform into a spherical shock wave. The initial strength of the spherical shock wave is determined from comparison with experimental results. Its subsequent propagation, and resulting pressure at the tube exit, is given by a numerical blast wave calculation. The model agrees reasonably well with other, limited, results. Finally, the model was used as the exit boundary condition for a one-dimensional calculation of PDE performance to obtain the thrust wall pressure for a hydrogen-air detonation in tubes of length to diameter ratio (L/D) of 4, and 10, as well as for the original, constant pressure boundary condition. The modified boundary condition had no performance impact for values of L/D > 10, and moderate impact for L/D = 4.
Wilson, Jack↗
Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program
The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems. All data gathered in this investigation are contained in the CD-ROM.
Asbury, Scott C.↗
Evaluation of an Active Clearance Control System Concept
Reducing blade tip clearances through active tip clearance control in the high pressure turbine can lead to significant reductions in emissions and specific fuel consumption as well as dramatic improvements in operating efficiency and increased service life. Current engines employ scheduled cooling of the outer case flanges to reduce high pressure turbine tip clearances during cruise conditions. These systems have relatively slow response and do not use clearance measurement, thereby forcing cold build clearances to set the minimum clearances at extreme operating conditions (e.g., takeoff, reburst) and not allowing cruise clearances to be minimized due to the possibility of throttle transients (e.g., step change in altitude). In an effort to improve upon current thermal methods, a first generation mechanically-actuated active clearance control (ACC) system has been designed and fabricated. The system utilizes independent actuators, a segmented shroud structure, and clearance measurement feedback to provide fast and precise active clearance control throughout engine operation. Ambient temperature performance tests of this first generation ACC system assessed individual seal component leakage rates and both static and dynamic overall system leakage rates. The ability of the nine electric stepper motors to control the position of the seal carriers in both open- and closed-loop control modes for single and multiple cycles was investigated. The ability of the system to follow simulated engine clearance transients in closed-loop mode showed the system was able to track clearances to within a tight tolerance (0.001 in. error).
Steinetz, Bruce M.↗
Evaluation of an Active Clearance Control System Concept
Reducing blade tip clearances through active tip clearance control in the high pressure turbine can lead to significant reductions in emissions and specific fuel consumption as well as dramatic improvements in operating efficiency and increased service life. Current engines employ scheduled cooling of the outer case flanges to reduce high pressure turbine tip clearances during cruise conditions. These systems have relatively slow response and do not use clearance measurement, thereby forcing cold build clearances to set the minimum clearances at extreme operating conditions (e.g., takeoff, reburst) and not allowing cruise clearances to be minimized due to the possibility of throttle transients (e.g., step change in altitude). In an effort to improve upon current thermal methods, a first generation mechanically-actuated active clearance control (ACC) system has been designed and fabricated. The system utilizes independent actuators, a segmented shroud structure, and clearance measurement feedback to provide fast and precise active clearance control throughout engine operation. Ambient temperature performance tests of this first generation ACC system assessed individual seal component leakage rates and both static and dynamic overall system leakage rates. The ability of the nine electric stepper motors to control the position of the seal carriers in both open- and closed-loop control modes for single and multiple cycles was investigated. The ability of the system to follow simulated engine clearance transients in closed-loop mode showed the system was able to track clearances to within a tight tolerance ( 0.001 in. error).
Steinetz, Bruce M.↗
The Transient Multi-Level method for Monte Carlo reactor statics calculations
The Transient Multi-Level (TML) method is applied to a time-dependent Monte Carlo transport solver to offload some of the computational burden of the expensive Monte Carlo solve to lower-order Coarse Mesh Finite Difference (CMFD) and Exact Point Kinetics Equations (EPKE) solvers via factorization of the neutron flux at the transport and CMFD levels using the Predictor Corrector Quasi-Static Method (PCQM). The Monte Carlo transient is solved by a modified fission source iteration scheme that introduces a single transient source bank. The method is implemented in the production-level Monte Carlo code, Shift, and verified with prescribed reactivity ramps from the two-dimensional version of the C5G7-TD reactor benchmark. The results show that, as compared to other quasi-static methods, the TML reduces the stochastic noise inherent to the transient Monte Carlo solver by factors of ~2 to 6 for various norm comparisons of the reactor power amplitude. Finally, the TML additionally reduces the number of Monte Carlo evaluations needed to simulate the transient, leading to roughly an order of magnitude improvement in CPU time relative to the standard PCQM for the problems tested.
22 GENERAL STUDIES OF NUCLEAR REACTORS↗
Static investigation of the circulation control wing/upper surface blowing concept applied to the quiet short haul research aircraft
Full scale static investigations were conducted on the Quiet Short Haul Research Aircraft (QSRA) to determine the thrust deflecting capabilities of the circulation control wing/upper surface blowing (CCW/USB) concept. This scheme, which combines favorable characteristics of both the A-6/CCW and QSRA, employs the flow entrainment properties of CCW to pneumatically deflect engine thrust in lieu of the mechanical USB flap system. Results show that the no moving parts blown system produced static thrust deflections in the range of 40 deg to 97 deg (depending on thrust level) with a CCW pressure of 208,900 Pa (30.3 psig). In addition, the ability to vary horizontal forces from thrust to drag while maintaining a constant vertical (or lift) value was demonstrated by varying the blowing pressure. The versatility of the CCW/USB system, if applied to a STOL aircraft, was confirmed, where rapid conversion from a high drag approach mode to a thrust recovering waveoff or takeoff configuration could be achieved by nearly instantaneous blowing pressure variation.
Eppel, J. C.↗
Integrating interface slicing into software engineering processes
Interface slicing is a tool which was developed to facilitate software engineering. As previously presented, it was described in terms of its techniques and mechanisms. The integration of interface slicing into specific software engineering activities is considered by discussing a number of potential applications of interface slicing. The applications discussed specifically address the problems, issues, or concerns raised in a previous project. Because a complete interface slicer is still under development, these applications must be phrased in future tenses. Nonetheless, the interface slicing techniques which were presented can be implemented using current compiler and static analysis technology. Whether implemented as a standalone tool or as a module in an integrated development or reverse engineering environment, they require analysis no more complex than that required for current system development environments. By contrast, conventional slicing is a methodology which, while showing much promise and intuitive appeal, has yet to be fully implemented in a production language environment despite 12 years of development.
Beck, Jon↗
BLASIM - A computational tool to assess ice impact damage on engine blades
A portable computer code called BLASIM is developed at NASA LeRC to assess the ice impact damage on aircraft engine blades. In addition to the ice impact analyses, the code is also capable of carrying out static, dynamic, resonance margin and flutter analyses. The blade can be solid, hollow, superhybrid or composite material. An optional preprocessor (input generator) is also developed to generate input to the code through interactive process. The blade geometry can be defined either by a series of airfoils at discrete input stations or by a finite element grid. The code employs a coarse fixed finite element mesh with triangular plate finite elements and has quick turnaround time. The ice piece is modeled as an equivalent spherical object and has the velocity opposite to that of the aircraft with direction parallel to the engine axis. For the local impact damage assessment, the impact force is considered as a distributed load acting over a region around the impact point and the average radial strain of the finite elements along the leading edge is taken as a measure of the local damage. To estimate the damage at the blade root, the impact is considered to be an impulse and a combined stress failure criteria is employed. Parametric studies for local and root ice impact damage, and post-impact dynamics are discussed for solid and composite blades.
Reddy, E. S.↗
Simulation of Flight-Type Engine Fan Noise in the NASA-Lewis 9X15 Anechoic Wind Tunnel
Flight type noise as contrasted to the usual ground static test noise exhibits substantial reductions in the time unsteadiness of tone noise, and in the mean level of tones calculated to be nonpropagating or cut-off. A model fan designed with cuttoff of the fundamental tone was acoustically tested in the anechoic wind tunnel under both static and tunnel flow conditions. The properties that characterize flight type noise were progressively simulated with increasing tunnel flow. The distinctly lobed directivity pattern of propagating rotor/stator interaction modes was also observed. Excess noise attributed to the ingestion of the flow disturbances that prevail near most static test facilities is substantially reduced with tunnel flow.
Heidmann, M. F.↗
STARS: A general-purpose finite element computer program for analysis of engineering structures
STARS (Structural Analysis Routines) is primarily an interactive, graphics-oriented, finite-element computer program for analyzing the static, stability, free vibration, and dynamic responses of damped and undamped structures, including rotating systems. The element library consists of one-dimensional (1-D) line elements, two-dimensional (2-D) triangular and quadrilateral shell elements, and three-dimensional (3-D) tetrahedral and hexahedral solid elements. These elements enable the solution of structural problems that include truss, beam, space frame, plane, plate, shell, and solid structures, or any combination thereof. Zero, finite, and interdependent deflection boundary conditions can be implemented by the program. The associated dynamic response analysis capability provides for initial deformation and velocity inputs, whereas the transient excitation may be either forces or accelerations. An effective in-core or out-of-core solution strategy is automatically employed by the program, depending on the size of the problem. Data input may be at random within a data set, and the program offers certain automatic data-generation features. Input data are formatted as an optimal combination of free and fixed formats. Interactive graphics capabilities enable convenient display of nodal deformations, mode shapes, and element stresses.
Gupta, K. K.↗
Propulsion systems noise technology
Turbofan engine noise research relevant to conventional aircraft is discussed. In the area of fan noise, static to flight noise differences were discussed and data were presented for two different ways of simulating flight behavior. Experimental results from a swept rotor fan design are presented which show that this concept has potential for reducing the multiple-pure-tone or buzz-saw noise related to the shock waves on a fan operating at supersonic tip speeds. Acoustic suppressor research objectives centered around the effect of the wave system generated by the fan stage that is the input to the treatment. A simplifying and unifying parameter, mode cutoff ratio was described. Results are presented which show that suppressor performance can be improved if the input wave is more precisely described. In jet noise, calculated results showing the potential noise reduction from the use of internal mixer nozzles rather than separate flow nozzles are presented.
Feiler, C. E.↗
Acoustic tests of duct-burning turbofan jet noise simulation: Comprehensive data report. Volume 1, section 1: Model scale data
Model scale data on a 12.2 m (40 ft) arc are presented which were obtained in the hot, static acoustic tests on eleven nozzle designs suitable for use on duct-burning turbofan engines.
Heck, P. H.↗