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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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NASA NTRS2008

Overview of NASA's Integrated Design and Engineering Analysis (IDEA)Environment

Historically, the design of subsonic and supersonic aircraft has been divided into separate technical disciplines (such as propulsion, aerodynamics and structures) each of which performs their design and analysis in relative isolation from others. This is possible in most cases either because the amount of interdisciplinary coupling is minimal or because the interactions can be treated as linear. The design of hypersonic airbreathing vehicles, like NASA s X-43, is quite the opposite. Such systems are dominated by strong non-linear interactions between disciplines. The design of these systems demands that a multi-disciplinary approach be taken. Furthermore, increased analytical fidelity at the conceptual design phase is highly desirable as many of the non-linearities are not captured by lower fidelity tools. Only when these systems are designed from a true multi-disciplinary perspective can the real performance benefits be achieved and complete vehicle systems be fielded. Toward this end, the Vehicle Analysis Branch at NASA Langley Research Center has been developing the Integrated Design & Engineering Analysis (IDEA) Environment. IDEA is a collaborative environment for parametrically modeling conceptual and preliminary launch vehicle configurations using the Adaptive Modeling Language (AML) as the underlying framework. The environment integrates geometry, configuration, propulsion, aerodynamics, aerothermodynamics, trajectory, closure and structural analysis into a generative, parametric, unified computational model where data is shared seamlessly between the different disciplines. Plans are also in place to incorporate life cycle analysis tools into the environment which will estimate vehicle operability, reliability and cost. IDEA is currently being funded by NASA s Hypersonics Project, a part of the Fundamental Aeronautics Program within the Aeronautics Research Mission Directorate. The environment is currently focused around a two-stage-to-orbit configuration with a turbine based combined cycle (TBCC) first stage and reusable rocket second stage. This paper provides an overview of the development of the IDEA environment, a description of the current status and detail of future plans.

Robinson, Jeffrey S.↗

NASA NTRS1973

Thermal control and structures approach for fluorinated propulsion.

FLOX and fluorine oxidizers require cryogenic storage at 86 K (-305 F) aboard a spacecraft, whereas their respective fuels MMH and hydrazine require room temperature storage. Propulsion configuration trade-off studies have been made for two potential classes of missions, an outer planet orbiter utilizing RTG's and an inner planet orbiter utilizing solar panels. The preliminary designs incorporate the fuel tank within the thermally controlled electronics compartment and expose the cryogenic tank to space. The cryogenic tank is isolated with foam, shadow shields and a titanium truss support. Its vernier temperature control is accomplished with electrical heaters.

Stimpson, L. D.↗

NASA NTRS1982

Main-belt asteroid exploration - Mission options for the 1990s

Mission configurations, propulsion systems, and target bodies for possible NASA asteroid exploration projects are examined. Noting that an announced delay in the development of a solar electric propulsion system has led to a consideration of chemical rocket systems, asteroid missions are grouped in terms of five potential areas for investigation, each successively further from the sun. The Shuttle-launched IUS is suggested as the prime candidate for boosting probes into trajectories for asteroid rendezvous with a number of the 3000 known asteroids. Planetary swingbys are mentioned as the only suitable method for satisfying the large energy requirements of the asteroid missions. Performance analyses are presented of the IUS 2-stage/Star-48 and Centaur vehicles, and sample missions to Fortuna, Anahita, and Urania in 1990 and further missions to the middle, outer, and Trojans asteroids are outlined.

Yen, C.-W. L.↗

NASA NTRS1982

Space Shuttle/high energy upper stage capabilities for the 1990's

Possible performance gains and cost reductions available through the evolution of succeedingly larger unmanned, and then manned, orbital transfer vehicles (OTV) as Shuttle upper stages are projected. Future missions could include delivery of 10,000 lb to GEO, planetary missions in the 2000-12,000 lb class, 30-42 ft payloads in the 5000-10,000 lb class, and manned and unmanned satellite servicing by the turn of the century. The vehicles could evolve from the Centaur F vehicle through stages of all-propulsive configurations to aerobraked, fully reusable vehicles. Reusability introduces cost savings and the ability to make plane changes. Furthermore, aerobraking will double the payload capability for round trip journeys to GEO, bringing costs down to $7000/lb.

Teixeira, C.↗

NASA NTRS1982

An engine trade study for a supersonic STOVL fighter-attack aircraft, volume 1

The best main engine for an advanced STOVL aircraft flight demonstrator was studied. The STOVL aircraft uses ejectors powered by engine bypass flow together with vectored core exhaust to achieve vertical thrust capability. Bypass flow and core flow are exhausted through separate nozzles during wingborne flight. Six near term turbofan engines were examined for suitability for this aircraft concept. Fan pressure ratio, thrust split between bypass and core flow, and total thrust level were used to compare engines. One of the six candidate engines was selected for the flight demonstrator configuration. Propulsion related to this aircraft concept was studied. A preliminary candidate for the aircraft reaction control system for hover attitude control was selected. A mathematical model of transfer of bypass thrust from ejectors to aft directed nozzle during the transition to wingborne flight was developed. An equation to predict ejector secondary air flow rate and ram drag is derived. Additional topics discussed include: nozzle area control, ejector to engine inlet reingestion, bypass/core thrust split variation, and gyroscopic behavior during hover.

Beard, B. B.↗

NASA NTRS1985

Readying technology for a super SST

A technology readiness evaluation is made of the configurational, propulsion system, avionics, and materials/manufacturing methods which could be integrated to produce an economically viable SST by the year 2000. Aerodynamically, promising results are noted in the cases of arrow-wing planforms and blended wing/body cross sections, as well as in vortex flaps. In structures, metal matrix composites with high specific strength at high temperatures and thermoplastically formed/diffusion-bonded titanium alloy sandwich structures are identified as being of primary importance. Performance improvement prospects are also noted for engine efficiencies and sonic boom overpressures. The various technologies are envisaged as the bases for a two-engine, two-crew cockpit, 250-seat arrow-winged SST.

Petersen, R. H.↗

NASA NTRS1988

Overview of the US/UK ASTOVL program

An account is given of progress to date in a US/UK advanced, supersonic flight-capable, powered-lift aircraft development collaborative effort formalized in January, 1986 by a memorandum of understanding (MOU). MOU-related work has investigated such ASTOVL propulsion configuration concepts as the remote augmented lift system, using a two-mode powerplant; augmenting ejectors, which generate a more benign ground environment; plenum chamber burning vectored thrust, which elaborates the geometry of the Harrier Pegasus powerplant; and tandem fans, which are also closely related to the Harrier concept but incorporate a truly variable-cycle engine.

Armstrong, Frank W.↗

NASA NTRS1988

Liquid rocket booster study. Volume 2, book 4, appendices 6-8: Reports of Rocketdyne, Pratt and Whitney, and TRW

For the pressure fed engines, detailed trade studies were conducted defining engine features such as thrust vector control methods, thrust chamber construction, etc. This was followed by engine design layouts and booster propulsion configuration layouts. For the pump fed engines parametric performance and weight data was generated for both O2/H2 and O2/RP-1 engines. Subsequent studies resulted in the selection of both LOX/RP-1 and O2/H2 propellants for the pump fed engines. More detailed analysis of the selected LOX/RP-1 and O2/H2 engines was conducted during the final phase of the study.

Source record↗

NASA NTRS1991

A comparison of chemical propulsion, nuclear thermal propulsion, and multimegawatt electric propulsion for Mars missions

Various propulsion systems are considered for a split-mission piloted exploration of Mars in terms of reducing total initial mass in low earth orbit (IMLEO) as well as trip time. Aerobraked nuclear thermal propulsion (NTP), multimegawatt (MMW) nuclear electric propulsion (NEP), and MMW solar electric propulsion (SEP) are discussed and compared to a baseline aerobraked chemical propulsion system. NTP offers low IMLEO, MMW NEP allows both low IMLEO and a short trip time, and both nuclear systems offer better mission characteristics than the chemical system. The MMW SEP is concluded to be less efficient in spite of a lower IMLEO because of the system's higher specific mass and nonconstant power production. It is recommended that MMW NEP and SEP systems be considered for application to Mars cargo missions. The NEP system is concluded to be the most effective propulsion configuration for piloted Mars missions and lunar base missions.

Frisbee, Robert H.↗

NASA NTRS2005

Technologies for Turbofan Noise Reduction

An overview presentation of NASA's engine noise research since 1992 is given for subsonic commercial aircraft applications. Highlights are included from the Advanced Subsonic Technology (AST) Noise Reduction Program and the Quiet Aircraft Technology (QAT) project with emphasis on engine source noise reduction. Noise reduction goals for 10 EPNdB by 207 and 20 EPNdB by 2022 are reviewed. Fan and jet noise technologies are highlighted from the AST program including higher bypass ratio propulsion, scarf inlets, forward-swept fans, swept/leaned stators, chevron nozzles, noise prediction methods, and active noise control for fans. Source diagnostic tests for fans and jets that have been completed over the past few years are presented showing how new flow measurement methods such as Particle Image Velocimetry (PIV) have played a key role in understanding turbulence, the noise generation process, and how to improve noise prediction methods. Tests focused on source decomposition have helped identify which engine components need further noise reduction. The role of Computational AeroAcoustics (CAA) for fan noise prediction is presented. Advanced noise reduction methods such as Hershel-Quincke tubes and trailing edge blowing for fan noise that are currently being pursued n the QAT program are also presented. Highlights are shown form engine validation and flight demonstrations that were done in the late 1990's with Pratt & Whitney on their PW4098 engine and Honeywell on their TFE-731-60 engine. Finally, future propulsion configurations currently being studied that show promise towards meeting NASA's long term goal of 20 dB noise reduction are shown including a Dual Fan Engine concept on a Blended Wing Body aircraft.

Huff, Dennis↗

NASA NTRS2012

Estimation of Uncertainties for a Supersonic Retro-Propulsion Model Validation Experiment in a Wind Tunnel

A high-quality model validation experiment was performed in the NASA Langley Research Center Unitary Plan Wind Tunnel to assess the predictive accuracy of computational fluid dynamics (CFD) models for a blunt-body supersonic retro-propulsion configuration at Mach numbers from 2.4 to 4.6. Static and fluctuating surface pressure data were acquired on a 5-inch-diameter test article with a forebody composed of a spherically-blunted, 70-degree half-angle cone and a cylindrical aft body. One non-powered configuration with a smooth outer mold line was tested as well as three different powered, forward-firing nozzle configurations: a centerline nozzle, three nozzles equally spaced around the forebody, and a combination with all four nozzles. A key objective of the experiment was the determination of experimental uncertainties from a range of sources such as random measurement error, flowfield non-uniformity, and model/instrumentation asymmetries. This paper discusses the design of the experiment towards capturing these uncertainties for the baseline non-powered configuration, the methodology utilized in quantifying the various sources of uncertainty, and examples of the uncertainties applied to non-powered and powered experimental results. The analysis showed that flowfield nonuniformity was the dominant contributor to the overall uncertainty a finding in agreement with other experiments that have quantified various sources of uncertainty.

Rhode, Matthew N.↗

NASA NTRS2009

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

Felder, James L.↗

NASA NTRS2018

Using Numerical Simulations to Explore Top-Mounted Propulsion on a Conceptual Commercial Supersonic Transport Aircraft

Reynolds Averaged Navier-Stokes (RANS) simulations were performed on an N+2 conceptual commercial supersonic transport aircraft to explore the possibility of relocating the engines from below the wings to above the wings in order to capitalize upon potential noise shielding benefits. The simulations focused on the feasibility of the top-mounted propulsion configuration in terms of inlet performance and flow separation around the nacelles at take-off conditions for both 0deg and 8deg angle of attack. The results showed overall good inlet performance and little separation around the nacelles. The results were comparable to the engines in their original underwing configuration.

Friedlander, David J.↗

NASA NTRS2019

Mars-GRAM Support for the Mars Ascent Vehicle

Mars sample return is a bold concept, which entails gathering a varied, scientifically-relevant collection of Martian rock core samples and bringing them to Earth for analysis. To support this endeavor, the Marshall Space Flight Center (MSFC) is developing the Mars Ascent Vehicle (MAV), which is responsible for getting the collected samples off the planet. The MAV Preliminary Architecture Assessment (PAA) study is designing two vehicle architectures based on different propulsion configurations: a two-stage solid-solid concept, and a hybrid concept. Given different thrust profiles for the two configurations, each concept uses a unique trajectory to reach the same orbit. In support of the PAA, The MSFC Natural Environments Branch (EV44) was asked to produce tables of atmospheric parameters along each of the two trajectories. The Mars Global Reference Atmospheric Model (Mars-GRAM) is an EV44 tool that is ideally suited for this analysis. Mars-GRAM will continue supporting MAV development in future design cycles.

Burns, K. L.↗

NASA NTRSDate not supplied

Results of the NASA Prediction Uncertainty Reduction Tech Challenge

In January 2021 the Advanced Air Vehicles Program approved a new Tech Challenge (TC) to be run out of the Commercial Supersonic Technology (CST) Project. The objective of the TC was to bring uncertainties in the empirical noise modeling for civilian supersonic aircraft into the same range as that of conventional aircraft. The TC goal statement was to “Produce data and demonstrate tools that reduce the uncertainty in predicting Landing & Takeoff Noise levels of supersonic-relevant designs by 5 EPNdB and are ready for use in studies to inform FAA rule-making.” To make the required improvements in noise prediction methods for supersonic aircraft, NASA decided to validate and use high-fidelity numerical simulations to acquire the needed noise data on supersonic propulsion configurations. High-fidelity component designs were developed, such as a two-stage propulsor behind a supersonic inlet designed by General Electric Aviation and variable area exhaust systems using input from recent contracts with GE and Rolls Royce. A noise database was generated as a function of geometric and flow parameters, providing corrections to the empirical noise models and added new input variables to describe the complexities created by the supersonic-specific design features. Statistically, the new models were able to predict the acoustic impact of supersonic-specific features, the reduction in uncertainty being reduced from the baseline 7.8 EPNdB at the beginning of the TC to 2.0 EPNdB at the end.

noise prediction↗

NASA NTRS2017

Structural Configuration Analysis of Advanced Flight Vehicle Concepts with Distributed Hybrid-Electric Propulsion

Structural configuration analysis of two advanced aircraft concepts with distributed hybrid-electric propulsion is presented. These concepts are characterized by multiple wing-mounted electric propulsors, which are powered by turbo-generators. Based on lessons learned from previous structural analysis of unconventional concepts, high-fidelity finite element models of the aircraft wing with embedded electric propulsors are developed. Although a hybrid-electric propulsion system has noise and emission benefits, it also adds electrical power system weights. Hence, efficient structural integration of the wing and propulsors is investigated for design improvement, structural analysis, and weight reduction. Wing structural weights of the two designs are compared with a baseline conventional transport aircraft wing for benefit assessment. In one design, the wing structural weight reduction partially compensates for the additional weight associated with the distributed electric propulsion system.

Mukhopadhyay, Vivek↗

NASA NTRS2020

Power and Propulsion Element for Gateway: Configuration and Data Management of the NASA Power and Propulsion Element MBSE Model(s)

The following will be presented: CM/DM implementation environment for MBSE, traditional construct of CM/MBSE, goal for power and propulsion element team for CM/MBSE, steps to bridge the gap between CM and MBSE, how we implemented the standard tenets of CM, realized efficiencies, encountered challenges, lessons learned, and current state, and forward/future work.

Parrott, Edith L.↗

NASA NTRS1984

Nuclear electric propulsion (NEP) spacecraft configuration study

The NEP spacecraft design is for the 100 kWe SP-100 nuclear reactor thermoelectric power supply. Thermoelectric conversion, one of several conversion options being considered for the SP-100, is discussed. Whereas many of the earlier NEP spacecraft designs have been based on highly integrated power and thruster subsystems and have resulted in subsystem requirements that depended largely on the design mission, the SP-100 nuclear power system is a general-purpose power supply. The configuration design must be compatible with this power subsystem design philosophy. The study also detects power and propulsion subsystem integration issues and related technology requirements. It is noted that the new NEP spacecraft design satisfies the most important subsystem integration requirements at the expense of some increase in vehicle inert mass relative to that used in previous studies.

Garrison, P. W.↗