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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 NTRS2010

NASA Innovation Fund 2010 Project Elastically Shaped Future Air Vehicle Concept

This report describes a study conducted in 2010 under the NASA Innovation Fund Award to develop innovative future air vehicle concepts. Aerodynamic optimization was performed to produce three different aircraft configuration concepts for low drag, namely drooped wing, inflected wing, and squashed fuselage. A novel wing shaping control concept is introduced. This concept describes a new capability of actively controlling wing shape in-flight to minimize drag. In addition, a novel flight control effector concept is developed to enable wing shaping control. This concept is called a variable camber continuous trailing edge flap that can reduce drag by as much as 50% over a conventional flap. In totality, the potential benefits of fuel savings offered by these concepts can be significant.

Nguyen, Nhan↗

NASA NTRSDate not supplied

Implementation Approach for an Electrified Aircraft Concept Vehicle in a Research Flight Simulator

•To describe the development of a dynamic model of an electrified propulsion integrated aircraft concept vehicle appropriate for implementation in a research flight simulator. •To bring up questions related to flight decks for electrified propulsion integrated aircraft. →Ultimate implementation of the model should be based on the procedure outlined here and more fully described in the paper →The flight deck requirements for a 2040 entry-into-service type vehicle with electrified propulsion need to be defined and implemented to the extent possible within the existing cockpit

flight simulation↗

NASA NTRS2001

Structures for the 3rd Generation Reusable Concept Vehicle

A major goal of NASA is to create an advance space transportation system that provides a safe, affordable highway through the air and into space. The long-term plans are to reduce the risk of crew loss to 1 in 1,000,000 missions and reduce the cost of Low-Earth Orbit by a factor of 100 from today's costs. A third generation reusable concept vehicle (RCV) was developed to assess technologies required to meet NASA's space access goals. The vehicle will launch from Cape Kennedy carrying a 25,000 lb. payload to the International Space Station (ISS). The system is an air breathing launch vehicle (ABLV) hypersonic lifting body with rockets and uses triple point hydrogen and liquid oxygen propellant. The focus of this paper is on the structural concepts and analysis methods used in developing the third generation reusable launch vehicle (RLV). Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented.

Hrinda, Glenn A.↗

NASA NTRS1993

Advanced transportation system study: Manned launch vehicle concepts for two way transportation system payloads to LEO. Program cost estimates document

This report describes Rockwell International's cost analysis results of manned launch vehicle concepts for two way transportation system payloads to low earth orbit during the basic and option 1 period of performance for contract NAS8-39207, advanced transportation system studies. Vehicles analyzed include the space shuttle, personnel launch system (PLS) with advanced launch system (ALS) and national launch system (NLS) boosters, foreign launch vehicles, NLS-2 derived launch vehicles, liquid rocket booster (LRB) derived launch vehicle, and cargo transfer and return vehicle (CTRV).

Duffy, James B.↗

NASA NTRS1981

Vehicle concepts and technology requirements for buoyant heavy-lift systems

Several buoyant-vehicle (airship) concepts proposed for short hauls of heavy payloads are described. Numerous studies identified operating cost and payload capacity advantages relative to existing or proposed heavy-lift helicopters for such vehicles. Applications involving payloads of from 15 tons up to 800 tons were identified. The buoyant quad-rotor concept is discussed in detail, including the history of its development, current estimates of performance and economics, currently perceived technology requirements, and recent research and technology development. It is concluded that the buoyant quad-rotor, and possibly other buoyant vehicle concepts, has the potential of satisfying the market for very heavy vertical lift but that additional research and technology development are necessary. Because of uncertainties in analytical prediction methods and small-scale experimental measurements, there is a strong need for large or full-scale experiments in ground test facilities and, ultimately, with a flight research vehicle.

Ardema, M. D.↗

NASA NTRS1988

Orbital transfer vehicle concept definition and systems analysis study. Volume 11: Study extension 2 results

Work conducted in the second extension of the Phase A Orbit Transfer Vehicle Concept Definition and Systems Analysis Study is summarized. Four major tasks were identified: (1) define an initial OTV program consistent with near term Civil Space Leadership Initiative missions; (2) develop program evolution to long term advanced missions; (3) investigate the implications of current STS safety policy on an Aft Cargo Carrier based OTV; and (4) expand the analysis of high entry velocity aeroassist. An increased emphasis on the breath of OTV applications was undertaken to show the need for the program on the basis of the expansion of the nation's capabilities in space.

Willcockson, W. H.↗

NASA NTRSDate not supplied

Enhancements, Verification, and VMS Integration of VTOL Concept Vehicle Simulation Models

Advanced Rotorcraft Technology (ART) and the NASA Ames Aeromechanics branch have jointly developed FLIGHTLAB simulation models for Advanced Air Mobility (AAM) VTOL concept vehicles. The overarching purpose of the simulation model development is to establish a set of well defined reference vehicles for FLIGHTLAB users and the rotorcraft community. The ongoing research effort and enhancement of these AAM simulation models to fulfill the role of quality reference vehicles is this paper’s focus. The content of this paper expands on the established characteristics of these AAM models in three primary areas. First, enhancement of the lift+cruise and tiltwing models with elastic airframe properties is discussed. The process of setting up the elastic airframe model in FLIGHTLAB, as well as the impacts on flight characteristics are explained. The introduction of the elastic airframe modeling allows these models to be used in flight dynamics, loads, and vibration analysis of the configuration designs. Next, linear model generation from the enhanced simulation model is covered. Confirming the validity of the linearized models is of importance, as these linear models are utilized for flight control design and tuning for these experimental configurations. For the final focus, the progress towards implementation of these models into the NASA Ames Vertical Motion Simulator (VMS) is described. This task seeks to demonstrate the procedures of integrating a FLIGHTLAB flight simulation in the VMS environment, test fully integrated simulation with communication between flight dynamics, control, and propulsion models, and explore the essential aspects of simulation model integration in a full flight simulator environment. This includes I/O definition, initialization, trim, flying, etc. By expanding the capabilities of the AAM simulation models, they continue to develop as valuable and approachable modeling references.

VMS↗

NASA NTRS2013

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three-stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and RP-1 on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 concept investigated potential impacts to affordability due to changes in gross liftoff weight and/or vehicle complexity. Results are discussed at a high level to understand the impact severity of certain sensitivities and how those trade studies conducted can either affect cost, performance, or both.

Beers, Benjamin R.↗

NASA NTRS2014

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three-stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current and future industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 concept investigated potential impacts to affordability due to changes in gross liftoff mass and/or vehicle complexity. Results are discussed at a high level to understand the impact severity of certain sensitivities and how those trade studies conducted can either affect cost, performance, or both.

Beers, Benjamin R.↗

NASA NTRS2014

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 configuration investigated potential impacts to affordability due to changes in gross liftoff weight and/or vehicle complexity. Results are discussed at a high level to understand the severity of certain sensitivities and how those trade studies conducted can either affect cost, performance or both.

Beers, Benjamin R.↗

NASA NTRSDate not supplied

Testing of Two Mars Powered Descent Vehicle Concepts in the Langley Unitary Plan Wind Tunnel

Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). Each model was designed to accommodate up to eight nozzles, with the HIAD model having five different nozzle configurations to investigate the effects of nozzle location, cant angle, and area ratio. The models were tested with high pressure air as the nozzle plume gas, and included the following instrumentation: high-speed video, discrete steady state and high-frequency pressure, pressure sensitive paint, and a new flow-through force and moment balance for the HIAD model. The high-speed imagery showed the overall expected growth of the shock layer to increasing thrust levels. The discrete stagnation pressure data on the HIAD model with four and eight blowing nozzles was sensitive to thrust coefficient with the nozzles canted at 0 degrees, whereas having nozzles canted outward 20 degrees and/or being located closer to the heatshield shoulder largely removed that sensitivity. The CobraMRV model stagnation pressure was more sensitive to tunnel Mach number and sideslip angle, due to the nozzle arrangement and plume interference. Pressure sensitive paint data quality was compromised by paint damage from frequent model changes, especially for the HIAD model. However, pressure coefficient data on the CobraMRV model showed the same trends as the discrete pressure measurements, as well as a sensitivity to non-zero sideslip angles. The aerodynamic force coefficients were derived from the pressure sensitive paint data. On the HIAD models, the primary force coefficient decreased with increasing thrust due to the nozzle plumes blocking flow to the heatshield area surrounding the nozzle exist. On the CobraMRV model, the force coefficient was relatively insensitive to thrust coefficient at the lower Mach number when sideslip angle was 0 degrees The force coefficient decreases with increasing thrust coefficient when the sideslip angle is 10 degrees. Balance data quality was negatively impacted by thermal drift issues that were not apparent in pre-test calibration measurements, thus preventing usable test data. Results and lessons learned will be used to take further technology development steps, including more advanced ground test techniques and flight testing.

Karl T Edquist↗

NASA NTRSDate not supplied

Testing of Two Mars Powered Descent Vehicle Concepts in the Langley Unitary Plan Wind Tunnel

Supersonic Retropropulsion↗

NASA NTRSDate not supplied

Computational Aerodynamics Study of the Lift+Cruise VTOL Concept Vehicle Components

This paper describes a building-block approach for high-fidelity computational fluid dynamic simulations of NASA’s Lift+Cruise (L+C) Vertical Take-off and Landing (VTOL) concept vehicle. The Reynolds-Averaged Navier-Stokes (RANS) equations are solved on overset structured grids using OVERFLOW. For these analyses, overset meshes are generated using recently developed automated meshing tools. A baseline study is initially performed on the vehicle fuselage with wings and tails. Vehicle components such as pylons and gears are then added individually and together to study incremental component aerodynamic effects. The results from this study are also used to demonstrate the capabilities of the automatic meshing tools for performing rapid computational analyses.

RVLT↗

NASA NTRSDate not supplied

Efficient Parametric Uncertainty Analysis of an Earth Entry Vehicle Concept Using Least Angle Regression

The objective of this work was to outline and apply an efficient and accurate parametric un-certainty propagation approach to the analysis of convective heating on an Earth entry vehicle concept. The described approach was based on Least Angle Regression used to solve a sparse and underdetermined linear system in the point-collocation non-intrusive polynomial chaos surrogate method. This approach involved an iterative process to computing the non-zero terms of the underlying polynomial chaos model using only enough samples to converge uncertainty interval predictions and Sobol index values based global nonlinear sensitivity estimates. The Earth entry vehicle was analyzed at three points along a representative trajectory for a Mars return mission. 329 sources of uncertainty were identified in the computational fluid dynamics model used to predict the forebody convective heating. These included uncertainty in flow field chemical rates, collision integrals, heats of formation, surface finite rate char model reaction rates, wall roughness height, and the turbulent Schmidt number. Results from this study showed that convective heating uncertainty as high as 50% of the nominal was predicted with only about 50 evaluations of the computational model. This was far fewer than would be required for a sampling-based approach or a full basis polynomial chaos model, which would have required over 50,000 samples. Additionally, results showed that over 90% of the total convective heating uncertainty was due to uncertainty in the N2catalytic rate on the surface, while the remainder of the uncertainty was attributed to the turbulent Schmidt number and the wall roughness uncertainties.

Thomas K West IV↗

NASA NTRSDate not supplied

Electrical System Trade Study for SUSAN Electrofan Concept Vehicle

The SUSAN Electrofan Aircraft concept is a subsonic regional jet transport aircraft that utilizes a single turbofan engine and Electrified Aircraft Propulsion (EAP) to enable propulsive and aerodynamic benefits to reduce fuel usage, emissions, and cost. This paper is a trade space exploration of the electrical system for the concept aircraft, including possible system architectures, and a review of the specific powers and efficiencies required for the electrical components to realize the EAP system.

Joseph M Haglage↗

NASA NTRS1985

Aeronautical technology 2000: A projection of advanced vehicle concepts

The Aeronautics and Space Engineering Board (ASEB) of the National Research Council conducted a Workshop on Aeronautical Technology: a Projection to the Year 2000 (Aerotech 2000 Workshop). The panels were asked to project advances in aeronautical technologies that could be available by the year 2000. As the workshop was drawing to a close, it became evident that a more comprehensive investigation of advanced air vehicle concepts than was possible in the limited time available at the workshop would be valuable. Thus, a special panel on vehicle applications was organized. In the course of two meetings, the panel identified and described representative types of aircraft judged possible with the workshop's technology projections. These representative aircraft types include: military aircraft; transport aircraft; rotorcraft; extremely high altitude aircraft; and transatmospheric aircraft. Improvements in performance, efficiency, and operational characteristics possible through the application of the workshop's year 2000 technology projections were discussed. The subgroups also identified the technologies considered essential and enhancing or supporting to achieve the projected aircraft improvements.

Source record↗

NASA NTRS1984

Aeronautical technology 2000 - A projection of advanced vehicle concepts

At the request of NASA and under the aegis of the National Research Council, representatives from industry, academic institutions and government have participated in a workshop to consider opportunities for the exploitation of aircraft technology in such fields as aerodynamics, materials, structures, guidance, navigation and control, human factors, propulsion, computers and data processing, and systems integration. Attention is given to the advanced vehicle concepts that have emerged for possible year-2000 implementation, which encompass such diverse aircraft types as supersonic transports, hypersonic airliners, missiles, and interceptors, transatmospheric vehicles, next-generation space shuttles, subsonic transports and attack aircraft, advanced helicopter, tilt-rotor VTOL configurations, and solar- and microwave beam-powered extremely high altitude aircraft.

Rosen, C. C., III↗

NASA NTRSDate not supplied

Computational Modeling of Two Mars Powered Descent Vehicle Concepts Tested in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not suitable for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flowfield predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at lower HIAD CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Supersonic Retropropulsion↗