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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Challenges of short-haul air transportation

Short-haul transportation systems are characterized by low average utilization, high peak but low average passenger demand, severe competition from other transportation modes, high fixed costs and high operating costs. Studies were conducted to analyze the operational requirements and market size for medium-density air transportation. The determination of the aircraft that will most efficiently meet the requirements of this market requires the careful evaluation of variable parameters such as engine cycles, payload, field length range and type of high-lift system. Low noise, low cost and thrust reversing capability are the basic criteria for the propulsion system. It is concluded that aircraft of less than 50-seat capacity cannot generate satisfactory profits operating in the medium-density market.

Galloway, T. L.↗
NASA NTRS1977

Quiet Clean Short-haul Experimental Engine (QCSEE): The aerodynamic and mechanical design of the QCSEE under-the-wing fan

The design, fabrication, and testing of two experimental high bypass geared turbofan engines and propulsion systems for short haul passenger aircraft are described. The aerodynamic and mechanical design of a variable pitch 1.34 pressure ratio fan for the under the wing (UTW) engine are included. The UTW fan was designed to permit rotation of the 18 composite fan blades into the reverse thrust mode of operation through both flat pitch and stall pitch directions.

Source record↗
NASA NTRS1985

Effects of propeller rotation direction on airplane interior noise levels

Interior noise measurements for upsweeping and downsweeping movement of the propeller blade tips past the fuselage were made on a twin-engine airplane and on two simplified fuselage models. Changes in interior noise levels of as much as 8 dB reversal of propeller rotation direction were measured for some configurations and test conditions.

Willis, C. M.↗
NASA NTRS1982

Equivalent damage: A critical assessment

Concepts in equivalent damage were evaluated to determine their applicability to the life prediction of hot path components of aircraft gas turbine engines. Equivalent damage was defined as being those effects which influence the crack initiation life-time beyond the damage that is measured in uniaxial, fully-reversed sinusoidal and isothermal experiments at low homologous temperatures. Three areas of equivalent damage were examined: mean stress, cumulative damage, and multiaxiality. For each area, a literature survey was conducted to aid in selecting the most appropriate theories. Where possible, data correlations were also used in the evaluation process. A set of criteria was developed for ranking the theories in each equivalent damage regime. These criteria considered aspects of engine utilization as well as the theoretical basis and correlative ability of each theory. In addition, consideration was given to the complex nature of the loading cycle at fatigue critical locations of hot path components; this loading includes non-proportional multiaxial stressing, combined temperature and strain fluctuations, and general creep-fatigue interactions. Through applications of selected equivalent damage theories to some suitable data sets it was found that there is insufficient data to allow specific recommendations of preferred theories for general applications. A series of experiments and areas of further investigations were identified.

Laflen, J. R.↗
NASA NTRS2015

NASA Engineering and Safety Center Technical Bulletin No. 15-01: Preventing Incorrect Installation of Polarized Capacitors

Even though NASA’s Lessons Learned Information System has an entry on the reverse installation of polarized capacitors, this error keeps reoccurring and is a “lesson not learned.” The purpose of this bulletin is to discuss some of the reasons this keeps happening as well as offer best practices to prevent this error in the future.

Preventing Incorrect Installation of Polarized Cap↗
NASA NTRS1994

A Revolutionary Lunar Space Transportation System Architecture Using Extraterrestrial Lox-augmented NTR Propulsion

The concept of a liquid oxygen (LOX)-augmented nuclear thermal rocket (NTR) engine is introduced, and its potential for revolutionizing lunar space transportation system (LTS) performance using extraterrestrial 'lunar-derived' liquid oxygen (LUNOX) is outlined. The LOX-augmented NTR (LANTR) represents the marriage of conventional liquid hydrogen (LH2)-cooled NTR and airbreathing engine technologies. The large divergent section of the NTR nozzle functions as an 'afterburner' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the NTR's choked sonic throat: 'scramjet propulsion in reverse.' By varying the oxygen-to-fuel mixture ratio (MR), the LANTR concept can provide variable thrust and specific impulse (Isp) capability with a LH2-cooled NTR operating at relatively constant power output. For example, at a MR = 3, the thrust per engine can be increased by a factor of 2.75 while the Isp decreases by only 30 percent. With this thrust augmentation option, smaller, 'easier to develop' NTR's become more acceptable from a mission performance standpoint (e.g., earth escape gravity losses are reduced and perigee propulsion requirements are eliminated). Hydrogen mass and volume is also reduced resulting in smaller space vehicles. An evolutionary NTR-based lunar architecture requiring only Shuttle C and/or 'in-line' shuttle-derived launch vehicles (SDV's) would operate initially in an 'expandable mode' with NTR lunar transfer vehicles (LTV's) delivering 80 percent more payload on piloted missions than their LOX/LH2 chemical propulsion counterparts. With the establishment of LUNOX production facilities on the lunar surface and 'fuel/oxidizer' depot in low lunar orbit (LLO), monopropellant NTR's would be outfitted with an oxygen propellant module, feed system, and afterburner nozzle for 'bipropellant' operation. The LANTR cislunar LTV now transitions to a reusable mode with smaller vehicle and payload doubling benefits on each piloted round trip mission. As the initial lunar outposts grow to centralized bases and settlements with a substantial permanent human presence, a LANTR-powered shuttle capable of 36 to 24 hour 'one-way' trip times to the moon and back becomes possible with initial mass in low earth orbit (IMLEO) requirements of approximately 160 to 240 metric tons, respectively.

Borowski, Stanley K.↗
NASA NTRS1928

Combustion time in the engine cylinder and its effect on engine performance

As part of a general program to study combustion in the engine cylinder and to correlate the phenomena of combustion with the observed performance of actual engines, this paper presents a sketchy outline of what may happen in the engine cylinder during the burning of a charge. It also suggests the type of information needed to supply the details of the picture and points out how combustion time and rate affect the performance of the engine. A theoretical concept of a flame front which is assumed to advance radially from the point of ignition is presented, and calculations based on the area and velocity of this flame and the density of the unburned gases are made to determine the mass rate of combustion. From this rate the mass which has been burned and the pressure at any instant during combustion are computed. This process is then reversed in an effort to determine actual rates of combustion and flame velocities from the pressures as recorded on indicator diagrams. The effects of different rates of combustion on engine performance are then discussed and the importance of proper spark advance is emphasized.

Marvin, Charles F , Jr↗
NASA NTRS1979

Quiet Clean Short-haul Experimental Engine (QCSEE); acoustic performance of a 50.8-cm (20 inch) diameter variable pitch fan and inlet, test results and analysis, volume 1

Tests were run both in forward and in reverse thrust modes with a bellmouth inlet, five accelerating inlets (one hard wall and four treated) with a design throat Mach number of 0.79 at the takeoff condition, and four low Mach inlets (one hard wall and three treated) with a design throat Mach number of 0.6 at the takeoff condition. Unsuppressed and suppressed inlet radiated noise levels were measured at conditions representative of QCSEE takeoff, approach, and reverse thrust operations. Measured aerodynamic performance of the accelerating inlet is also included. The test objectives, facility, configurations, are described as well as the data analysis, results, and comparisons.

Bilwakesh, K. R.↗
NASA NTRS1997

Investigation of the Reverse Water Gas Shift Reaction for Production of Oxygen From Mars Atmospheric Carbon Dioxide

The first phase of the research includes a comprehensive analytical study examining the potential applications for engineering subsystems and mission strategies made possible by such RWGS based subsystems, and will include an actual experimental demonstration and performance characterization of a full-scale brassboard RWGS working unit. By the time of this presentation the laboratory demonstration unit will not yet be operational but we will present the results of our analytical studies to date and plans for the ongoing work.

Meyer, Tom↗
NASA NTRS1983

Small Gas Turbine Combustor Primary Zone Study

A development process is described which consists of design, fabrication, and preliminary test evaluations of three approaches to internal aerodynamic primary zone flow patterns: (1) conventional double vortex swirl stabilization; (2) reverse flow swirl stabilization; and (3) large single vortex flow system. Each concept incorporates special design features aimed at extending the performance capability of the small engine combustor. Since inherent geometry of these combustors result in small combustion zone height and high surface area to volume ratio, design features focus on internal aerodynamics, fuel placement, and advanced cooling. The combustors are evaluated on a full scale annular combustor rig. A correlation of the primary zone performance with the overall performance is accomplished using three intrusion type gas sampling probes located at the exit of the primary zone section. Empirical and numerical methods are used for designing and predicting the performance of the three combustor concepts and their subsequent modifications. The calibration of analytical procedures with actual test results permits an updating of the analytical design techniques applicable to small reverse flow annular combustors.

Sullivan, R. E.↗
NASA NTRS1983

Clean catalytic combustor program

A combustor program was conducted to evolve and to identify the technology needed for, and to establish the credibility of, using combustors with catalytic reactors in modern high-pressure-ratio aircraft turbine engines. Two selected catalytic combustor concepts were designed, fabricated, and evaluated. The combustors were sized for use in the NASA/General Electric Energy Efficient Engine (E3). One of the combustor designs was a basic parallel-staged double-annular combustor. The second design was also a parallel-staged combustor but employed reverse flow cannular catalytic reactors. Subcomponent tests of fuel injection systems and of catalytic reactors for use in the combustion system were also conducted. Very low-level pollutant emissions and excellent combustor performance were achieved. However, it was obvious from these tests that extensive development of fuel/air preparation systems and considerable advancement in the steady-state operating temperature capability of catalytic reactor materials will be required prior to the consideration of catalytic combustion systems for use in high-pressure-ratio aircraft turbine engines.

Ekstedt, E. E.↗
NASA NTRS1982

Small gas turbine combustor primary zone development

Designers of small gas turbine engines prefer a close-coupled compressor to turbine shafting arrangement, which in some designs necessitates the use of a small reverse-flow annular combustor. A design methodology for obtaining the maximum performance potential of these combustors is necessary. This paper describes an approach to optimize the design process and gain insight into primary zone performance through interactive theoretical analyses and experimental tests. Three candidate combustor designs are described which address the performance limiting problem areas associated with small annular combustors. Design methodology centers around understanding and controlling primary zone aerodynamics and the interaction of the distributed fuel with internal airflow patterns. Complete three-dimensional flow field analytical performance prediction procedures are presented and results compared with performance and emission measurements described by probes located at the exit of the primary zone. The effective use of analytical performance prediction methods in the design process is demonstrated.

Sullivan, R. E.↗
NASA NTRS2001

MHD Energy Bypass Scramjet Engine

Revolutionary rather than evolutionary changes in propulsion systems are most likely to decrease cost of space transportation and to provide a global range capability. Hypersonic air-breathing propulsion is a revolutionary propulsion system. The performance of scramjet engines can be improved by the AJAX energy management concept. A magneto-hydro-dynamics (MHD) generator controls the flow and extracts flow energy in the engine inlet and a MHD accelerator downstream of the combustor accelerates the nozzle flow. A progress report toward developing the MHD technology is presented herein. Recent theoretical efforts are reviewed and ongoing experimental efforts are discussed. The latter efforts also include an ongoing collaboration between NASA, the US Air Force Research Laboratory, US industry, and Russian scientific organizations. Two of the critical technologies, the ionization of the air and the MHD accelerator, are briefly discussed. Examples of limiting the combustor entrance Mach number to a low supersonic value with a MHD energy bypass scheme are presented, demonstrating an improvement in scramjet performance. The results for a simplified design of an aerospace plane show that the specific impulse of the MHD-bypass system is better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Equilibrium ionization and non-equilibrium ionization are discussed. The thermodynamic condition of air at the entrance of the engine inlet determines the method of ionization. The required external power for non-equilibrium ionization is computed. There have been many experiments in which electrical power generation has successfully been achieved by magneto-hydrodynamic (MHD) means. However, relatively few experiments have been made to date for the reverse case of achieving gas acceleration by the MHD means. An experiment in a shock tunnel is described in which MHD acceleration is investigated experimentally. MHD has several potential aerospace applications. The first is to improve the performance of hypersonic air-breathing engines for space launch and cruise vehicles. The second is to improve the performance of a high enthalpy wind tunnel. The third is to control a hypersonic vehicle. With such applications in mind, theoretical and experiments are being conducted at the NASA Ames Research Center to develop the MHD technology.

Mehta, Unmeel B.↗
NASA NTRS1974

Anthropometric changes and fluid shifts

Several observations of body size, shape, posture, and configuration were made to document changes resulting from direct effects of weightlessness during the Skylab 4 mission. After the crewmen were placed in orbit, a number of anatomical and anthropometric changes occurred including a straightening of the thoracolumbar spine, a general decrease in truncal girth, and an increase in height. By the time of the earliest in-flight measurement on mission day 3, all crewmen had lost more than two liters of extravascular fluid from the calf and thigh. The puffy facies, the bird legs effect, the engorgement of upper body veins, and the reduced volume of lower body veins were all documented with photographs. Center-of-mass measurements confirmed a fluid shift cephalad. This shift remained throughout the mission until recovery, when a sharp reversal occurred; a major portion of the reversal was completed in a few hours. The anatomical changes are of considerable scientific interest and of import to the human factors design engineer, but the shifts of blood and extravascular fluid are of more consequence. It is hypothesized that the driving force for the fluid shift is the intrinsic and unopposed lower limb elasticity that forces venous blood and then other fluid cephalad.

Thornton, W. E.↗
NASA NTRS1983

Dilution Jet Experiments in Compact Combustor Configurations

The effects of cooling jets on the velocity and temperature fields in a compact reverse flow combustion are discussed. The work is motivated by the need to limit the temperature of past-combustion gases in jet engines to values within the endurance capabilities of turbine blades. The application requires not only that the temperature be kept sufficiently low but also that a suitably tailored temperature profile be provided at the combustor exit, with higher temperatures generally permissible at the blade tip than at the blade root because of higher centrifugal loads at the root. A single jet. As spacing is reduced, jet penetration is also reduced, and the cooling jets tend to remain close to the wall from which they are injected. Results suggest that cooling and temperature distribution tailoring can be accomplished without injecting cooling jets upstream of the turn, and thus it appears that combustors can be made smaller than current designs.

Greber, I.↗
NASA NTRS1973

Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model

Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 percent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 percent for an 85 percent efficient fan stage.

Wesoky, H. L.↗
NASA NTRS1973

Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model

Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 per cent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 per cent for an 85 per cent efficient fan stage. For comparison, a similar 1.55 pressure ratio fan system would have a propulsive efficiency of 62 per cent.

Wesoky, H. L.↗