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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

The Space Shuttle - Orbital flight phase begins

The results of the first orbital flight test of the Space Shuttle are summarized and the future operations are discussed. It is mentioned that increased payload lift capability will be obtained through a series of planned steps which have already been initiated in the program. These include the use of the main engines at higher thrust levels, lower weight External Tank and Solid Rocket Boosters (SRB's), higher thrust SRB's, and removal of certain test equipment from the Orbiter. The economics of the Space Shuttle to the users and the future uses of the Space Shuttle are also discussed.

Day, L. E.↗
NASA NTRS1986

Modification of integrated partial payload lifting assembly

The Integrated Partial Payload Lifting Assembly (IPPLA) is currently used to transport and load experimental payloads into the cargo bay of the Space Shuttle. It is unable to carry the astronaut/passenger tunnel without a structural modification. The purpose of this design is to create a removalbe modification that will allow the IPPLA to lift and carry the passenger tunnel. Modifications evaluated were full-length insert beams which would extend through the existing strongback arms. These beam proposals were eliminated because of high cost and weight. Other proposals evaluated were attachments of cantilever beams to the existing strongback areas. The cantilever proposals reduced cost and weight compared to the full-length modifications. A third method evaluated was to simply make modifications to one side of the IPPLA therefore reducing the materials of the cantilever proposals by 40 percent. The design of the modification selected was completed with two channel beams jointly welded to a centered steel plate. The extension arm modification is inserted into the existing strongback channel beams and bolted into place. Two extension arms are added to one side of the IPPLA to provide the extra length needed to accommodate the passenger tunnel. The center counterbalance will then be offset about 20 inches to center gravity and therefore maintain horizontal status. The extension arm modification was selected because of minimum cost, low weight, and minimal installation time.

Groah, Melodie↗
NASA NTRS1999

Correlation Between Geometric Similarity of Ice Shapes and the Resulting Aerodynamic Performance Degradation: A Preliminary Investigation Using WIND

Aerodynamic performance calculations were performed using WIND on ten experimental ice shapes and the corresponding ten ice shapes predicted by LEWICE 2.0. The resulting data for lift coefficient and drag coefficient are presented. The difference in aerodynamic results between the experimental ice shapes and the LEWICE ice shapes were compared to the quantitative difference in ice shape geometry presented in an earlier report. Correlations were generated to determine the geometric features which have the most effect on performance degradation. Results show that maximum lift and stall angle can be correlated to the upper horn angle and the leading edge minimum thickness. Drag coefficient can be correlated to the upper horn angle and the frequency-weighted average of the Fourier coefficients. Pitching moment correlated with the upper horn angle and to a much lesser extent to the upper and lower horn thicknesses.

Wright, William B.↗
NASA NTRS2010

Amplitude Scaling of Active Separation Control

Three existing and two new excitation magnitude scaling options for active separation control at Reynolds numbers below one Million. The physical background for the scaling options was discussed and their relevance was evaluated using two different sets of experimental data. For F+ approx. 1, 2D excitation: a) The traditional VR and C(mu) - do not scale the data. b) Only the Re*C(mu) is valid. This conclusion is also limited for positive lift increment.. For F+ > 10, 3D excitation, the Re corrected C(mu), the St corrected velocity ratio and the vorticity flux coefficient, all scale the amplitudes equally well. Therefore, the Reynolds weighted C(mu) is the preferred choice, relevant to both excitation modes. Incidence also considered, using Ue from local Cp.

Stalnov, Oksana↗
NASA NTRS1968

Air Bearing Lift Pad (ABLP)

Typical air bearings float on air films of only a few thousandths of an inch and so will only operate above very smooth, even surfaces. For the mechanical simulation of space, the small drag of the bladder type air pads is much more than can be coped with, and the practicality of large floor areas being machined for precision air bearings is nonexistent. To enable operation above surfaces that undulate slightly or feature cracks and discontinuities, an ABLP has been developed. It consists of a rigid pad beneath which an inflatable bladder is mounted. The bladder is inflated with air which then escapes through passages into a cavity in the center of the bladder to produce the lifting energy. As the air escapes about the perimeter of the bladder, a certain degree of balance and equilibrium is imparted to the pad as it is able to move a limited weight across slightly uneven surfaces.

Dane, Dan H.↗
NASA NTRS2022

Low-Speed Performance Enhancement using Localized Active Flow Control: Program Overview and Summary (1/4)

The Boeing team executed a NASA task order, titled “Low Speed Performance Enhancement using Localized Active Flow Control”, under the BAART framework NNL16AA04B, contract number 80LARC20F0082. The project was executed from Sept. 2020 through April 2022 with the objective to explore localized active flow control (AFC) concepts on a representative commercial aircraft. A detailed literature review solidified the choice of AFC concepts to be explored, including AFC over a deflected aileron, in the leading edge region of the wing, and in the nacelle/pylon/wing region. All three concepts were investigated with numerical tools. The results were used in an integration study to assess the net benefits over the lifetime of a commercial aircraft. The aerodynamic studies concluded that AFC applied over a deflected aileron yields significant net L/D improvements (incl. the penalty of AFC system requirements) of up to potentially 5% L/D. Depending on the configuration, AFC applied in the leading edge region showed improvements in L/D of up to ~1.5%, as well as opportunities for maximum lift and increased lift at fixed angle of attack. AFC applied to the nacelle/pylon/wing region delivered 1.5% increase in L/D during take-off and ~4% increase in CLmax during landing. After considering assumed onboard sources, aspects of system integration and weight penalties, these aerodynamic improvements translate to: ~ 0.5% block fuel reduction potential for AFC over the aileron ~ 0.1% block fuel reduction potential for AFC in the LE slat region ~ 0.2% block fuel reduction potential for AFC in the nacelle/pylon/wing region considering take-off scenarios (potentially larger benefit may be realized considering the landing scenarios which haven’t been fully analyzed yet) The final report is comprised of four separate documents. The current document provides an overview and technical background. The second document details the numerical studies and the third document summarizes the integration and aircraft performance assessment. The fourth document focuses on the Common Research Model, specifically how AFC performs over a deflected aileron and how the NASA 10% scale wind tunnel model may be modified to implement an AFC equipped aileron. Each report offers significantly more detailed summaries and suggestions for future work.

Active flow control↗
NASA NTRS1975

A study of design trade (OFFS) using a computer model

The interaction between the efficiency of the structural design and the cost of the structure used was studied. It is shown that future effort is best directed at producing a low cost structure of medium efficiency, but with the ability to withstand normal service wear. The trade-off between aerodynamic drag and structure weight in selecting a length to diameter ratio for the hull is evaluated along with the implications of power plan type and fuel cost on the economics of the airship. The choice of lifting gas is considered.

Coughlin, S.↗
NASA NTRS1973

QCSEE task 2: Engine and installation preliminary design

High-bypass turbofan engines with features required for commercial short haul powered lift transports were designed. Two engines were configured for each of the externally blown flap installations, under-the-wing and over-the-wing. Estimates of installed and uninstalled performance, noise, and weight were defined for each propulsion system.

Neitzel, R. E.↗
NASA NTRS1984

Improving the efficiency of smaller transport aircraft

The efficiency of small transport aircraft can be improved through the adaptation of high altitude turbine engines, and that flights reaching altitudes of 40,000, 60,000, and 80,000 feet can show savings in both flight time and fuel consumption even for trips as short as 500 miles. Studies for a 40-passenger high altitude transport are presented. An increase in structural weight due to larger wing areas, larger engines, and larger engine frontal areas would make the ratio of gross weight to payload look less favorable, but the efficiency of the plane in passenger miles per gallon would increase with altitude. It is also suggested that supercritical airfoils be designed to achieve higher lift coefficients and speeds. A reduction of reduced drag through the use of horizontal or vertical wing tip extensions is also discussed.

Jones, R. T.↗
NASA NTRS1993

High-altitude balloon-launched aircraft - A piloted simulation study

A real-time piloted simulation at the NASA Dryden Flight Research Facility was used to study the feasibility of launching a research aircraft from a high-altitude balloon. In the study, the simulated aircraft was launched in a nosedown attitude at zero airspeed from 110,000 ft. After launch, the pilot flew the aircraft through a near-maximum-lift pullout and then through a zoom climb to a trimmed, l-g flight condition at the test altitude. The study included parametric variations to measure the effects of launch altitude, gross weight, Mach number limit, and parachute size on the test altitude attained. The aerodynamic model of the simulated aircraft was based on flight test results, low Reynolds number windtunnel tests, and computational models; the model included significant Mach number and Reynolds number effects at high altitude. A small parachute was included in the simulation to limit Mach number during the pullout to avoid adverse transonic effects and their resultant energy losses. A small rocket motor was included in the simulation and was investigated for boosting the aircraft to a higher test altitude. In the study, a test altitude of approximately 95,000 ft was attained without rocket boost, and a test altitude in excess of 100,000 ft was attained using small rocket boost.

Murray, James E.↗
NASA NTRS1994

Fast Track NTR Systems Assessment for NASA's First Lunar Outpost Scenario

Integrated systems and mission study results are presented which quantify the rationale and benefits for developing and using nuclear thermal rocket (NTR) technology for returning humans to the moon in the early 2000's. At present, the Exploration Program Office (ExPO) is considering chemical propulsion for its 'First Lunar Outpost' (FLO) mission, and NTR propulsion for the more demanding Mars missions to follow. The use of an NTR-based lunar transfer stage, capable of evolving to Mars mission applications, could result in an accelerated schedule, reduced cost approach to moon/Mars exploration. Lunar mission applications would also provide valuable operational experience and serve as a 'proving ground' for NTR engine and stage technologies. In terms of performance benefits, studies indicate that an expendable NTR stage powered by two 50 klbf engines can deliver approximately 96 metric tons (t) to trans-lunar injection (TLI) conditions for an initial mass in low earth orbit (IMLEO) of approximately 199 t compared to 250 t for a cryogenic chemical TLI stage. The NTR stage liquid hydrogen (LH2) tank has a 10 m diameter, 14.8 m length, and 68 t LH2 capacity. The NTR utilizes a 'graphite' fuel form consisting of coated UC2 particles in a graphite substrate, and has a specific impulse capability of approximately 870 s, and an engine thrust-to-weight ratio of approximately 4.8. The NTR stage and its piloted FLO lander has a total length of approximately 38 m and can be launched by a single Saturn V-derived heavy lift launch vehicle (HLLV) in the 200 to 250 t-class range. The paper summarizes NASA's First Lunar Outpost scenario, describes characteristics for representative engine/stage configurations, and examines the impact on engine selection and vehicle design resulting from a consideration of alternative NTR fuel forms and lunar mission profiles.

Borowski, Stanley K.↗
NASA NTRS2005

Beam-Steerable Flat-Panel Reflector Antenna

Many space applications require a high-gain antenna that can be easily deployable in space. Currently, the most common high-gain antenna for space-born applications is an umbrella-type reflector antenna that can be folded while being lifted to the Earth orbit. There have been a number of issues to be resolved for this type of antenna. The reflecting surface of a fine wire mesh has to be light in weight and flexible while opening up once in orbit. Also the mesh must be a good conductor at the operating frequency. In this paper, we propose a different type of high-gain antenna for easy space deployment. The proposed antenna is similar to reflector antennas except the curved main reflector is replaced by a flat reconfigurable surface for easy packing and deployment in space. Moreover it is possible to steer the beam without moving the entire antenna system.

Lee, Choon Sae↗
NASA NTRS1963

LATERAL-RANGE AND HYPERSONIC LIFT-DRAG-RATIO REQUIREMENTS FOR EFFICIENT FERRY SERVICE FROM A NEAR-EARTH MANNED SPACE STATION

The operation of any long-term manned space station will require some type of ferry vehicle to transport men and equipment to and from the station with regularity and reliability. Such a vehicle, designed for entry at near-orbital speeds, could also be useful in the return from any deeper space mission if either an earth-orbit rendezvous terminal maneuver or a maneuver combining atmospheric braking and a near-earth parking orbit is used. This study was undertaken to determine the class of vehicle which could be most efficiently used as a ferry vehicle between a near-earth space station and the earth. One measure of this efficiency is the ability of the vehicle to reach pre-chosen landing sites with some prescribed frequency. In considering this frequency of return it is necessary to consider not only the normal mode of operation in which only infrequent returns are scheduled at desirable times, but also operation under various degrees of emergency, which dictate quick or even immediate return to earth. In extreme emergencies, when immediate return to earth is necessary, choice of landing site becomes impractical. In most cases, however, although it might be required to abandon the station quickly, the ferry vehicle could remain in orbit for some time before initiating reentry in order to land at a prechosen site. The allowable delay time in orbit would be determined primarily by the capabilities of the ferry life-support system. This paper will examine the geometry of the ferry ranging problem, that is, the lateral ranges required to reach chosen landing sites from various near-earth orbits, and will investigate and compare several means of achieving these ranges. The particular case considered is that of returning from a space station which is in a circular orbit at an altitude of 200 statute miles, but the results obtained are not sensitive to orbit altitude for orbits within a few hundred miles of the surface. From this orbit, the vehicle will retro and reenter at very close to satellite velocity. The downrange problem can be handled by proper timing of the retrofiring , and the desired lateral range can be achieved by aerodynamics, space propulsion to change orbit plane, atmospheric propulsion, or combinations of these methods. The relative cost in terms of weight of using these different methods to achieve lateral range will be discussed.

Hypersonic vehicle↗
NASA NTRS2005

Blended-Wing-Body (BWB) Fuselage Structural Design for Weight Reduction

Structural analysis and design of efficient pressurized fuselage configurations for the advanced Blended-Wing-Body (BWB) flight vehicle is a challenging problem. Unlike a conventional cylindrical pressurized fuselage, stress level in a box type BWB fuselage is an order of magnitude higher, because internal pressure primarily results in bending stress instead of skin-membrane stress. In addition, resulting deformation of aerodynamic surface could significantly affect performance advantages provided by lifting body. The pressurized composite conformal multi-lobe tanks of X-33 type space vehicle also suffered from similar problem. In the earlier BWB design studies, Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS); Vaulted shell Honeycomb Core (VLHC) and Flat sandwich shell Honeycomb Core (FLHC) concepts were studied. The flat and vaulted ribbed shell concepts were found most efficient. In a recent study, a set of composite sandwich panel and cross-ribbed panel were analyzed. Optimal values of rib and skin thickness, rib spacing, and panel depth were obtained for minimal weight under stress and buckling constraints. In addition, a set of efficient multi-bubble fuselage (MBF) configuration concept was developed. The special geometric configuration of this concept allows for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls, while the outer-ribbed shell prevents buckling due to external resultant compressive loads. The initial results from these approximate finite element analyses indicate progressively lower maximum stresses and deflections compared to the earlier study. However, a relative comparison of the FEM weight per unit floor area of the segment unit indicates that the unit weights are still relatively higher that the conventional B777 type cylindrical or A380 type elliptic fuselage design. Due to the manufacturing concern associated with multi-bubble fuselage, a Y braced box-type fuselage alternative with special resin-film injected (RFI) stitched carbon composite with foam-core was designed by Boeing under a NASA research contract for the 480 passenger version. It is shown that this configuration can be improved to a modified multi-bubble fuselage which has better stress distribution, for same material and dimension.

Mukhopadhyay, V.↗
NASA NTRS1981

Computer program to predict aircraft noise levels

Methods developed at the NASA Lewis Research Center for predicting the noise contributions from various aircraft noise sources were programmed to predict aircraft noise levels either in flight or in ground tests. The noise sources include fan inlet and exhaust, jet, flap (for powered lift), core (combustor), turbine, and airframe. Noise propagation corrections are available for atmospheric attenuation, ground reflections, extra ground attenuation, and shielding. Outputs can include spectra, overall sound pressure level, perceived noise level, tone-weighted perceived noise level, and effective perceived noise level at locations specified by the user. Footprint contour coordinates and approximate footprint areas can also be calculated. Inputs and outputs can be in either System International or U.S. customary units. The subroutines for each noise source and propagation correction are described. A complete listing is given.

Clark, B. J.↗
NASA NTRS1995

Numerical Methodology For Full-Body TPS Sizing And Optimization For Access-To-Space Vehicles

This paper presents details of the methodology and numerical procedures developed at NASA Ames for full-body TEXAS sizing and optimization for Access to Space vehicle concepts. The core of the procedures is a robust implicit solver for one dimensional transient heat conduction in reusable multilayer TEXAS stackups. The solver includes an arbitrary number of material layers, contact resistances between materials, temperature and pressure dependent material and surface properties, numerous boundary-condition options, and self-adaptive time stepping. The solver is coupled with the Access-to-Space material database of 23 candidate TPS and structural materials and a thermal-environment database obtained from trajectory-based fullbody Navier-Stokes computations of the external flowfield. The thermal environment and material response are coupled through the use of T-type heat transfer coefficients. TIPS sizing and weight optimization are performed at every surface point on the vehicle based on sizing constraints which include material temperature limits, maximum backwall temperature, and cumulative interior heat flux. Typical results are presented for a lifting body concept with 10000 surface points, which required 35 minutes to compute on an SGI Indigo 2.

Milos, Frank S.↗
NASA NTRS2013

Development of the Circulation Control Flow Scheme Used in the NTF Semi-Span FAST-MAC Model

The application of a circulation control system for high Reynolds numbers was experimentally validated with the Fundamental Aerodynamic Subsonic Transonic Modular Active Control semi-span model in the NASA Langley National Transonic Facility. This model utilized four independent flow paths to modify the lift and thrust performance of a representative advanced transport type of wing. The design of the internal flow paths highlights the challenges associated with high Reynolds number testing in a cryogenic pressurized wind tunnel. Weight flow boundaries for the air delivery system were identified at mildly cryogenic conditions ranging from 0.1 to 10 lbm/sec. Results from the test verified system performance and identified solutions associated with the weight-flow metering system that are linked to internal perforated plates used to achieve flow uniformity at the jet exit.

Jones, Gregory S.↗
NASA NTRS2017

Evaluation of a Variable Thickness Hybrid Composite Bull Gear

For several years, NASA Glenn Research Center and the U.S. Army Research Laboratory have been investigating hybrid (composite/steel) gear technology for use in vertical lift drive systems. The hybrid gear concept replaces the structural portion of a gear between the shaft and the gear rim with a lightweight carbon fiber composite, in an effort to reduce the overall weight of a gear and increase the drive system power density. Past research includes both small-scale and large-scale hybrid gear concepts, all of which have a constant composite thickness throughout. The design described in this paper is of a variable thickness, such that the composite is thickest at the inner diameter and this thickness is gradually reduced toward the outer diameter. The resulting "stair stepped" design stems from dropping plies of the braided carbon fiber prepreg composite fabric gradually with increased radius. Additionally, the interlock pattern at the inner metallic adapter was adjusted slightly from previous designs to obtain a better stress distribution on the inner metallic adapter. The manufactured variable thickness web was tested both in static torsion tests and operationally in a relevant gearbox environment. The results of these experiments will be presented and compared to a baseline steel configuration.

Gears; Drive Systems; Rotorcraft Propulsion;↗