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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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At least 559 records · Page 31
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Influence of noise requirements on STOL propulsion system designs

The severity of proposed noise goals for STOL systems has resulted in a new design approach for aircraft propulsion systems. It has become necessary to consider the influence of the noise goal on the design of engine components, engine systems, and the integrated nacelle, separately and collectively, from the onset of the design effort. This integrated system design approach is required in order to effect an optimization of the propulsion and aircraft system. Results from extensive design studies and pertinent test programs are presented which show the effect of noise specifications on component and system design, and the trade offs possible of noise versus configuration and performance. The design optimization process of propulsion systems for powered lift systems is presented beginning with the component level and proceeding through to the final integrated propulsion system. Designs are presented which are capable of meeting future STOL noise regulations and the performance, installation and economic penalties are assessed as a function of noise level.

Rulis, R. J.↗
NASA NTRS1983

Scale model performance test investigation of mixed flow exhaust systems for an energy efficient engine /E3/ propulsion system

As part of the NASA Energy Efficient Engine program, scale-model performance tests of a mixed flow exhaust system were conducted. The tests were used to evaluate the performance of exhaust system mixers for high-bypass, mixed-flow turbofan engines. The tests indicated that: (1) mixer penetration has the most significant affect on both mixing effectiveness and mixer pressure loss; (2) mixing/tailpipe length improves mixing effectiveness; (3) gap reduction between the mixer and centerbody increases high mixing effectiveness; (4) mixer cross-sectional shape influences mixing effectiveness; (5) lobe number affects mixing degree; and (6) mixer aerodynamic pressure losses are a function of secondary flows inherent to the lobed mixer concept.

Kuchar, A. P.↗
NASA NTRS1976

Reverse-thrust technology for variable-pitch fan propulsion systems

Tests conducted to develop the technology necessary to meet the unique reverse-thrust performance requirements of a variable pitch fan propulsion system are discussed. The losses and distortion associated with the air entering the fan and core compressor from the rear of the engine, the direction of fan blade pitch rotation for best reverse-thrust aeroacoustic performance, and engine response and operating characteristics during forward- to reverse-thrust transients are among the factors studied. The test results of several scale fan models as well as a full-size variable pitch fan engine are summarized. Results show the following: a flared exhaust nozzle makes a good reverse-thrust inlet; acceptable core inlet duct recovery and distortion levels in reverse flow were demonstrated; adequate thrust levels were achieved; forward- to reverse-thrust response time achieved was better than the goal; thrust and noise levels strongly favor reverse through feather pitch; and finally, flight-type inlets make the establishment of reverse flow more difficult.

Sagerser, D. A.↗
NASA NTRS1977

Space Shuttle Plume and Plume Impingement Study

The extent of the influence of the propulsion system exhaust plumes on the vehicle performance and control characteristics is a complex function of vehicle geometry, propulsion system geometry, engine operating conditions and vehicle flight trajectory were investigated. Analytical support of the plume technology test program was directed at the two latter problem areas: (1) definition of the full-scale exhaust plume characteristics, (2) application of appropriate similarity parameters; and (3) analysis of wind tunnel test data. Verification of the two-phase plume and plume impingement models was directed toward the definition of the full-scale exhaust plume characteristics and the separation motor impingement problem.

Tevepaugh, J. A.↗
NASA NTRS1974

Influence of noise requirements on STOL propulsion system designs

The severity of proposed noise goals for STOL systems has resulted in a new design approach for aircraft propulsion systems. It has become necessary to consider the influence of the noise goal on the design of engine components, engine systems, and the integrated nacelle, separately and collectively, from the onset of the design effort. This integrated system design approach is required in order to effect an optimization of the propulsion and aircraft system. Results from extensive design studies and pertinent test programs are presented which show the effect of noise specifications on component and system design, and the trade-offs possible of noise versus configuration and performance. The design optimization process of propulsion systems for powered lift systems is presented beginning with the component level and proceeding through to the final integrated propulsion system. Designs are presented which are capable of meeting future STOL noise regulations and the performance, installation and economic penalties are assessed as a function of noise level.

Rulis, R. J.↗
NASA NTRS1973

Influence of noise requirements on STOL propulsion system designs

The severity of proposed noise goals for STOL systems has resulted in a new design approach for aircraft propulsion systems. It has become necessary to consider the influence of the noise goal on the design of engine components, engine systems, and the integrated nacelle, separately and collectively, from the onset of the design effort. This integrated system design approach is required in order to effect an optimization of the propulsion and aircraft system. Results from extensive design studies and pertinent test programs are presented which show the effect of noise specifications on component and system design, and the trade-offs possible of noise versus configuration and performance. The design optimization process of propulsion systems for powered lift systems is presented beginning with the component level and proceeding through to the final integrated propulsion system. Designs are presented which are capable of meeting future STOL noise regulations and the performance, installation and economic penalties are assessed as a function of noise level.

Rulis, R. J.↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Thruster Control Mode Design and Flight Experience

National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, designed, built, tested, and launched the Lunar Reconnaissance Orbiter (LRO) from Cape Canaveral Air Force Station on June 18, 2009. The LRO spacecraft is the first operational spacecraft designed to support NASA s return to the Moon, as part of the Vision for Space Exploration. LRO was launched aboard an Atlas V 401 launch vehicle into a direct insertion trajectory to the Moon. Twenty-four hours after separation the propulsion system was used to perform a mid-course correction maneuver. Four days after the mid-course correction a series of propulsion maneuvers were executed to insert LRO into its commissioning orbit. The commission period lasted eighty days and this followed by a second set of thruster maneuvers that inserted LRO into its mission orbit. To date, the spacecraft has been gathering invaluable data in support of human s future return to the moon. The LRO Attitude Control Systems (ACS) contains two thruster based control modes: Delta-H and Delta-V. The design of the two controllers are similar in that they are both used for 3-axis control of the spacecraft with the Delta-H controller used for momentum management and the Delta-V controller used for orbit adjust and maintenance maneuvers. In addition to the nominal purpose of the thruster modes, the Delta-H controller also has the added capability of performing a large angle slew maneuver. A suite of ACS components are used by the thruster based control modes, for both initialization and control. For initialization purposes, a star tracker or the Kalman Filter solution is used for providing attitude knowledge and upon entrance into the thruster based control modes attitude knowledge is provided via rate propagation using a inertial reference unit (IRU). Rate information for the controller is also supplied by the IRU. Three-axis control of the spacecraft in the thruster modes is provided by eight 5-lbf class attitude control thrusters configured in two sets of four thrusters for redundancy purposes. Four additional 20-lbf class thrusters configured in two sets of two thrusters are used for Lunar Orbit Insertion maneuvers. The propulsion system is one the few systems on-board the LRO spacecraft that has built in redundancy. The Delta-H controller consists of a Proportional-Derivative (PD) controller with a structural filter on the thrusters and a Proportional controller on the reaction wheels. The PD control that employs the thrusters is used for attitude and rate control. The Proportional controller on the reaction wheels is used for commanding the wheels to a new momentum state. The ground commands used for the Delta-H controller are the system momentum vector, reaction wheel momentum, maximum expected command time, and which set of attitude control thrusters to use. The ability to command both the system momentum vector and reaction wheel momentum in the Delta-H controller provides both a capability and an additional source of operator error. Large angle slews via the Delta-H controller is achievable via this commands because these commands are used for the exit mode criteria. Setting these commands to non-consistent values prevents the mode from exiting nominally.

Hsu, Oscar C.↗
NASA NTRS1981

Effect of time-dependent flight loads on turbofan engine performance deterioration

An analytic evaluation of the dynamic effects of two flight load conditions of the JT9D-7/747 propulsion system is conducted. Predicted performance changes associated with a once-per-flight vertical gust and a typical revenue service landing are calculated. The predicted dynamic load effects on thrust specific fuel consumption are found to be negligible. The results indicate that the quasi-steady state approach to flight loads modeling is adequate to investigate the factors important to the deterioration process.

Lewis, B. L.↗
NASA NTRS1992

Development of NASA/DOE NTP System Performance Models

A critical enabling technology in the evolutionary development of Nuclear Thermal Propulsion (NTP) is the ability to predict the system performance under a variety of operating conditions. The ability to predict the system performance is critical for mission analysis and for control subsystem testing, as well as for the modeling of various failure modes. Performance must be accurately predicted during steady-state and transient operation, such as start-up, shut-down and after-cooling. The development and application of verified and validated system models has the potential to reduce testing, cost and time required for the technology to again reach flight-ready status. An integrated NASA/DOE team was formed in late 1991 to develop and implement a strategy for modeling NTP systems. It is the intent of the interagency team to develop several levels of computer programs, which vary in detail, to simulate NTP systems based on either prismatic, particle or advanced fuel forms. This paper presents an overview of the models under development by the interagency team. In addition, the status of the development and validation efforts for the Level 1 steady-state parametric model is discussed.

Walton, James T.↗
NASA NTRS1974

Cruise performance of an isolated 1.15 pressure ratio turbofan propulsion system simulator at Mach numbers from 0.6 to 0.85

An isolated 1.15 pressure ratio turbofan engine simulator was tested at Mach numbers from 0.6 to 0.85. At Mach 0.75 the net propulsive force of the fan and nacelle (excluding core thrust) was 73 percent of the ideal fan net thrust. Internal losses amounted to 7 percent, and external drag amounted to 20 percent of the ideal fan net thrust. External pressure and friction drag were about equal. The propulsive efficiency with a 90 percent efficient fan would have been 63 percent. For the aerodynamic characteristics of the nacelle that was tested, increasing the fan pressure ratio to approximately 1.35 would have resulted in a maximum propulsive efficiency of 67 percent.

Steffen, F. W.↗
NASA NTRS1979

Characteristics of primary electric propulsion systems

The use of advanced electric propulsion systems will provide cost and performance benefits for future energetic space missions. A methodology to predict the characteristics of advanced electric propulsion systems was developed and programmed for computer calculations to allow evaluation of a broad set of technology and mission assumptions. The impact on overall thrust system characteristics was assessed for variations of propellant type, total accelerating voltage, thruster area, specific impulse, and power system approach. The data may be used both to provide direction to technology emphasis and allow for preliminary estimates of electric propulsion system properties for a wide variety of application.

Byers, D. C.↗
NASA NTRS2017

Annoyance to Noise Produced by a Distributed Electric Propulsion High-Lift System

A psychoacoustic test was performed using simulated sounds from a distributed electric propulsion aircraft concept to help understand factors associated with human annoyance. A design space spanning the number of high-lift leading edge propellers and their relative operating speeds, inclusive of time varying effects associated with motor controller error and atmospheric turbulence, was considered. It was found that the mean annoyance response varies in a statistically significant manner with the number of propellers and with the inclusion of time varying effects, but does not differ significantly with the relative RPM between propellers. An annoyance model was developed, inclusive of confidence intervals, using the noise metrics of loudness, roughness, and tonality as predictors.

Rizzi, Stephen A.↗
NASA NTRS2012

Quasi 1D Modeling of Mixed Compression Supersonic Inlets

The AeroServoElasticity task under the NASA Supersonics Project is developing dynamic models of the propulsion system and the vehicle in order to conduct research for integrated vehicle dynamic performance. As part of this effort, a nonlinear quasi 1-dimensional model of the 2-dimensional bifurcated mixed compression supersonic inlet is being developed. The model utilizes computational fluid dynamics for both the supersonic and subsonic diffusers. The oblique shocks are modeled utilizing compressible flow equations. This model also implements variable geometry required to control the normal shock position. The model is flexible and can also be utilized to simulate other mixed compression supersonic inlet designs. The model was validated both in time and in the frequency domain against the legacy LArge Perturbation INlet code, which has been previously verified using test data. This legacy code written in FORTRAN is quite extensive and complex in terms of the amount of software and number of subroutines. Further, the legacy code is not suitable for closed loop feedback controls design, and the simulation environment is not amenable to systems integration. Therefore, a solution is to develop an innovative, more simplified, mixed compression inlet model with the same steady state and dynamic performance as the legacy code that also can be used for controls design. The new nonlinear dynamic model is implemented in MATLAB Simulink. This environment allows easier development of linear models for controls design for shock positioning. The new model is also well suited for integration with a propulsion system model to study inlet/propulsion system performance, and integration with an aero-servo-elastic system model to study integrated vehicle ride quality, vehicle stability, and efficiency.

Kopasakis, George↗
NASA NTRS1989

Hybrid propulsion technology program: Phase 1, volume 1

A number of booster propulsion system concepts are being considered for the next generation of manned and unmanned space launch vehicles. The one propulsion system concept that has potential for reducing costs with increased safety, reliability, and performance is hybrid propulsion (HP). A HP system may be thought of as a liquid propulsion system with solid fuel or a solid propulsion system with a liquid oxidizer. The liquid propulsion features that are most attractive are the higher specific impulse, clean exhaust, separated propellants, and oxidizer loading just prior to launch. The most attractive solid propulsion features includes low life cycle costs, no rotating machinery, compact size, and a robust case. In addition, a HP system has a robust LO2 tank; provides thrust control for ignition, to alleviate flight loads, and for thrust termination; and uses an inert grain that is not sensitive to anomalies such as cracks, voids, and separations. The object is to develop the technology to enable the application of HP to manned and unmanned space launch vehicles. This program will identify the necessary technology, acquire that technology, and demonstrate that technology. This volume is the executive summary.

Schuler, A. L.↗
NASA NTRS1973

Study of advanced techniques for determining the long term performance of components

The application of existing and new technology to the problem of determining the long-term performance capability of liquid rocket propulsion feed systems is discussed. The long term performance of metal to metal valve seats in a liquid propellant fuel system is stressed. The approaches taken in conducting the analysis are: (1) advancing the technology of characterizing components through the development of new or more sensitive techniques and (2) improving the understanding of the physical of degradation.

Source record↗
NASA NTRS2017

Personal Rotorcraft Design and Performance with Electric Hybridization

Recent and projected improvements for more or all-electric aviation propulsion systems can enable greater personal mobility, while also reducing environmental impact (noise and emissions). However, all-electric energy storage capability is significantly less than present, hydrocarbon-fueled systems. A system study was performed exploring design and performance assuming hybrid propulsion ranging from traditional hydrocarbon-fueled cycles (gasoline Otto and diesel) to all-electric systems using electric motors generators, with batteries for energy storage and load leveling. Study vehicles were a conventional, single-main rotor (SMR) helicopter and an advanced vertical takeoff and landing (VTOL) aircraft. Vehicle capability was limited to two or three people (including pilot or crew); the design range for the VTOL aircraft was set to 150 miles (about one hour total flight). Search and rescue (SAR), loiter, and cruise-dominated missions were chosen to illustrate each vehicle and degree of hybrid propulsion strengths and weaknesses. The traditional, SMR helicopter is a hover-optimized design; electric hybridization was performed assuming a parallel hybrid approach by varying degree of hybridization. Many of the helicopter hybrid propulsion combinations have some mission capabilities that might be effective for short range or on-demand mobility missions. However, even for 30 year technology electrical components, all hybrid propulsion systems studied result in less available fuel, lower maximum range, and reduced hover and loiter duration than the baseline vehicle. Results for the VTOL aircraft were more encouraging. Series hybrid combinations reflective of near-term systems could improve range and loiter duration by 30. Advanced, higher performing series hybrid combinations could double or almost triple the VTOL aircrafts range and loiter duration. Additional details on the study assumptions and work performed are given, as well as suggestions for future study effort.

electric motors↗
NASA NTRSDate not supplied

Solar Array Performance Modeling for NASA’s Artemis Missions

NASA’s Artemis exploration campaign includes multiple elements that utilize photovoltaic power generation. Artemis mission profiles feature a variety of events that can limit solar array generation and stress electrical power system (EPS) performance, including eclipses, propulsive or navigational maneuvers which constrain the positioning of the solar arrays, and spiral trajectories with significant radiation degradation. These features make predicting the performance of the power system particularly important and requires accurate modeling of power generation by the solar arrays under a variety of conditions. One of the tools used to predict spacecraft EPS performance is the System Power Analysis for Capability Evaluation (SPACE) model developed at NASA Glenn Research Center. SPACE is used by NASA to model EPS performance for the Orion crew transport vehicle and the Power and Propulsion Element (PPE) of the Gateway space station. This presentation will provide an overview of how SPACE calculates solar array performance, including degradation factors considered and environmental conditions that drive the solar array designs. Additionally, Orion array performance predictions generated by SPACE will be compared to flight data collected during the Artemis I mission.

Photovoltaic↗
NASA NTRS2005

Solar Electric and Chemical Propulsion for a Titan Mission

Systems analyses were performed for a Titan Explorer Mission characterized by Earth-Saturn transfer stages using solar electric power generation and propulsion systems for primary interplanetary propulsion, and chemical propulsion for capture at Titan. An examination of a range of system factors was performed to determine their effect on the payload delivery capability to Titan. The effect of varying launch vehicle type, solar array power level, ion thruster number, specific impulse, trip time, and Titan capture stage chemical propellant choice was investigated. The major purpose of the study was to demonstrate the efficacy of applying advanced ion propulsion system technologies like NASA's Evolutionary Xenon Thruster (NEXT), coupled with state-of-the-art (SOA) and advanced chemical technologies to a Flagship class mission. This study demonstrated that a NASA Design Reference Mission (DRM) payload of 406 kg could be successfully delivered to Titan using the baseline advanced ion propulsion system in conjunction with SOA chemical propulsion for Titan capture. In addition, the SEPS/Chemical system of this study is compared to an all- chemical NASA DRM mission. Results showed that the NEXT- based SEPS/Chemical system was able to deliver the required payload to Titan in 5 years less transfer time and on a smaller launch vehicle than the SOA chemical option.

Cupples, Michael↗