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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program

The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems. All data gathered in this investigation are contained in the CD-ROM.

Asbury, Scott C.↗
NASA NTRS1981

Design predictions for noise control in the cryogenic National Transonic Facility

Noise control in the National Transonic Facility - a cryogenic wind tunnel - has been examined in terms of acoustical design criteria, drive-fan noise and exhaust system noise. A duct lining with two layers of perforated sheeting and a gas-filled honeycomb core was selected for attenuating drive-fan noise. With the exception of attenuation peaks, attenuation of the lining was found to experimentally agree with predicted values at 20 C air temperatures. Exhaust system noise will be attenuated with a large muffler used in conjunction with a 6.1-m high acoustical enclosure. Fan noise from the fan-ejector system will be attenuated by fan silencers and the acoustic enclosure.

Lassiter, W. S.↗
NASA NTRS1980

Experimental evaluation of exhaust mixers for an Energy Efficient Engine

Static scale model tests were conducted to evaluate exhaust system mixers for a high bypass ratio engine as part of the NASA sponsored Energy Efficient program. Gross thrust coefficients were measured for a series of mixer configurations which included variations in the number of mixer lobes, tailpipe length, mixer penetration, and length. All of these parameters have a significant impact on exhaust system performance. In addition, flow visualization pictures and pressure/temperature traverses were obtained for selected configurations. Parametric performance trends are discussed and the results considered relative to the Energy Efficient Engine program goals.

Kozlowski, H.↗
NASA NTRS2009

Broadband Shock Noise in Internally-Mixed Dual-Stream Jets

Broadband shock noise (BBSN) has been studied in some detail in single-flow jets and recently in dual-stream jets with separate flow exhaust systems. Shock noise is of great concern in these latter cases because of the noise created for the aircraft cabin by the underexpanded nozzle flow at cruise. Another case where shock noise is of concern is in the case of future supersonic aircraft that are expected to have bypass ratios small enough to justify internally mixed exhaust systems, and whose mission will push cycles to the point of imperfectly expanded flows. Dual-stream jets with internally mixed plume have some simplifying aspects relative to the separate flow jets, having a single shock structure given by the common nozzle pressure. This is used to separate the contribution of the turbulent shear layer to the broadband shock noise. Shock structure is held constant while the geometry and strength of the inner and merged shear layers are varying by changing splitter area ratio and core stream temperature. Flow and noise measurements are presented which document the efforts at separating the contribution of the inner shear layer to the broadband shock noise.

Bridges, James E.↗
NASA NTRS1979

High temperature UF6 RF plasma experiments applicable to uranium plasma core reactors

An investigation was conducted using a 1.2 MW RF induction heater facility to aid in developing the technology necessary for designing a self critical fissioning uranium plasma core reactor. Pure, high temperature uranium hexafluoride (UF6) was injected into an argon fluid mechanically confined, steady state, RF heated plasma while employing different exhaust systems and diagnostic techniques to simulate and investigate some potential characteristics of uranium plasma core nuclear reactors. The development of techniques and equipment for fluid mechanical confinement of RF heated uranium plasmas with a high density of uranium vapor within the plasma, while simultaneously minimizing deposition of uranium and uranium compounds on the test chamber peripheral wall, endwall surfaces, and primary exhaust ducts, is discussed. The material tests and handling techniques suitable for use with high temperature, high pressure, gaseous UF6 are described and the development of complementary diagnostic instrumentation and measurement techniques to characterize the uranium plasma, effluent exhaust gases, and residue deposited on the test chamber and exhaust system components is reported.

Roman, W. C.↗
NASA NTRS1984

Static jet noise test results of four 0.35 scale-model QCGAT mixer nozzles

As part of the NASA Quiet Clean General Aviation Turbofan (QCGAT) engine mixer-nozzle exhaust system program, static jet exhaust noise was recorded at microphone angles of 45 to 155 deg relative to the nozzle inlet for a conventional profile coaxial nozzle and three 12-lobed coaxial mixer nozzles. Both flows in all four nozzles are internally mixed before being discharged from a single exhaust nozzle. The conventional profile coaxial nozzle jet noise is compared to the current NASA Lewis coaxial jet noise prediction and after applying an adjustment to the predicted levels based on the ratio of the kinetic energy of the primary and secondary flows, the prediction is within a standard deviation of 0.9 dB of the measured data. The mass average (mixed flow) prediction is also compared to the noise data for the three mixer nozzles with a reasonably good fit after applying another kinetic energy ratio adjustment (standard deviation of 0.7 to 1.5 dB with the measured data). The tests included conditions for the full-scale engine at takeoff (T.O.), cutback (86% T.O.) and approach (67% T.O.).

Groesbeck, D. E.↗
NASA NTRS2002

Inerting Aircraft Fuel Systems Using Exhaust Gases

Our purpose in this proposal was to determine the feasibility of using carbon dioxide, possibly obtained from aircraft exhaust gases as a substance to inert the fuel contained in fuel tanks aboard aircraft. To do this, we decided to look at the effects carbon dioxide has upon commercial Jet-A aircraft fuel. In particular, we looked at the solubility of CO2 in Jet-A fuel, the pumpability of CO2-saturated Jet-A fuel, the flashpoint of Jet-A fuel under various mixtures of air and CO2, the static outgassing of CO2-Saturated Jet-A fuel and the dynamic outgassing of Jet-A fuel during pumping of Jet-A fuel.

Hehemann, David G.↗
NASA NTRS2008

System for Removing Pollutants from Incinerator Exhaust

A system for removing pollutants -- primarily sulfur dioxide and mixed oxides of nitrogen (NOx) -- from incinerator exhaust has been demonstrated. The system is also designed secondarily to remove particles, hydrocarbons, and CO. The system is intended for use in an enclosed environment, for which a prior NOx-and-SO2-removal system designed for industrial settings would not be suitable.

Wickham, David t.↗
NASA NTRS1991

An experimental comparison of nonswirling and swirling flow in a circular-to-rectangular transition duct

Circular-to-rectangular transition ducts are used as exhaust system components of aircraft with rectangular exhaust nozzles. Often, the incoming flow of these transition ducts includes a swirling velocity component remaining from the gas turbine engine. Previous transition duct studies have either not included inlet swirl or when inlet swirl was considered, only overall performance parameters were evaluated. Circular-to-rectangular transition duct flows with and without inlet swirl were explored in order to understand the effect of inlet swirl on the transition duct flow field and to provide detailed duct flow data for comparison with numerical code predictions. A method was devised to create a swirling, solid body rotational flow with minimal associated disturbances. Coefficients based on velocities and total and static pressures measured incross stream planes at four axial locations within the transition duct, along with surface static pressure measurements and surface oil film visualization, are presented for both nonswirling and swirling incoming flow. In both cases the inlet centerline Mach number was 0.35. The Reynolds number based on the inlet centerline velocity and duct inlet diameter was 1,547,000 for nonswirling and 1,366,000 for swirling flow. The maximum swirl angle was 15.6 deg. Two pair of counter-rotating side wall vortices appeared in the duct flow without inlet swirl. These vortices were absent in the swirling incoming flow cases.

Reichert, B. A.↗
NASA NTRSDate not supplied

Dew Point Simulation and Analysis of the Heat Melt Compactor Gas Effluent

The effluent gases and water vapor of the Trash Compaction and Processing System (TCPS) can be vented through the Vacuum Exhaust System (VES) if requirements are met. One of the constraints in this venting scenario is the dew point of the exhaust gases disposed into the VES must be less than 15.5 °C. With simulations using Aspen Plus and the effluent gas sample results of the Heat Melt Compactor (HMC), flash calculations were conducted in the modeling study to calculate the feed temperature and dew point at fixed pressures. Saturated vapor curves were produced and provided a preliminary result on optimal feed conditions that satisfy the dew point and vapor-phase only requirements after flashing across the throttle valve towards the VES.

trash↗
NASA NTRS1981

Advanced Gas Turbine (AGT) powertrain system development for automotive applications

Preliminary layouts were made for the exhaust system, air induction system, and battery installation. Points of interference were identified and resolved by altering either the vehicle or engine designs. An engine general arrangement evolved to meet the vehicle engine compartment constraints while minimizing the duct pressure losses and the heat rejection. A power transfer system (between gasifier and power turbines) was developed to maintain nearly constant temperatures throughout the entire range of engine operation. An advanced four speed automatic transmission was selected to be used with the engine. Performance calculations show improvements in component efficiencies and an increase in fuel economy. A single stage centrifugal compressor design was completed and released for procurement. Gasifier turbine, power turbine, combustor, generator, secondary systems, materials, controls, and transmission development are reported.

Source record↗
NASA NTRS1974

Surgical suite environmental control system

Theoretical and experimental work for a systems analysis approach to the problem of surgical suit exhaust systems centered on evaluation of halothane absorbing filters. An activated charcoal-alumina-charcoal combination proved to be the best filter for eliminating halothane through multilayer absorption of gas molecules.

Higginbotham, E. J.↗
NASA NTRS1999

NASA Glenn 1-by 1-Foot Supersonic Wind Tunnel User Manual

This manual describes the NASA Glenn Research Center's 1 - by 1 -Foot Supersonic Wind Tunnel and provides information for customers who wish to conduct experiments in this facility. Tunnel performance envelopes of total pressure, total temperature, and dynamic pressure as a function of test section Mach number are presented. For each Mach number, maps are presented of Reynolds number per foot as a function of the total air temperature at the test section inlet for constant total air pressure at the inlet. General support systems-such as the service air, combustion air, altitude exhaust system, auxiliary bleed system, model hydraulic system, schlieren system, model pressure-sensitive paint, and laser sheet system are discussed. In addition, instrumentation and data processing, acquisition systems are described, pretest meeting formats and schedules are outlined, and customer responsibilities and personnel safety are addressed.

Seablom, Kirk D.↗
NASA NTRS2003

Summary of Fluidic Thrust Vectoring Research Conducted at NASA Langley Research Center

Interest in low-observable aircraft and in lowering an aircraft's exhaust system weight sparked decades of research for fixed geometry exhaust nozzles. The desire for such integrated exhaust nozzles was the catalyst for new fluidic control techniques; including throat area control, expansion control, and thrust-vector angle control. This paper summarizes a variety of fluidic thrust vectoring concepts that have been tested both experimentally and computationally at NASA Langley Research Center. The nozzle concepts are divided into three categories according to the method used for fluidic thrust vectoring: the shock vector control method, the throat shifting method, and the counterflow method. This paper explains the thrust vectoring mechanism for each fluidic method, provides examples of configurations tested for each method, and discusses the advantages and disadvantages of each method.

Deere, Karen A.↗
NASA NTRS2015

Oxide_Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project

Rolls-Royce North American Technologies, Inc. (LibertyWorksLW) began considering the development of CMC exhaust forced mixers in 2008, as a means of obtaining reduced weight and hotter operating temperature capability, while minimizing shape distortion during operation, which would improve mixing efficiency and reduce fuel burn. Increased component durability, enhanced ability to fabricate complex-shaped components, and engine noise reduction are other potential advantages of CMC mixers (compared to metallic mixers). In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project. ERA subtasks, including those focused on CMC components, were formulated with the goal of maturing technology from proof of concept validation (TRL 3) to a systemsubsystem or prototype demonstration in a relevant environment (TRL 6). In April 2010, the NASA Glenn Research Center (GRC) and LibertyWorks jointly initiated a CMC Exhaust System Validation Program within the ERA Project, teaming on CMC exhaust mixer development for subsonic jet engines capable of operating with increased performance. Our initial focus was on designing, fabricating, and characterizing the thrust and acoustic performance of a roughly quarter-scale 16-lobe oxide oxide CMC mixer and tail cone along with a conventional low bypass exhaust nozzle. Support Services, LLC (Allendale, MI) and ATK COI Ceramics, Inc. (COIC, in San Diego, CA) supported the design of a subscale nozzle assembly that consisted of an oxide oxide CMC mixer and center body, with each component mounted on a metallic attachment ring. That design was based upon the operating conditions a mixer would experience in a turbofan engine. Validation of the aerodynamic and acoustic performance of the subscale mixer via testing and the achievement of TRL 4 encouraged the NASALWCOIC team to move to the next phase where a full scale CMC mixer sized for a RR AE3007 engine and a compatible attachment flange were designed, followed by CMC component fabrication by COIC, and vibration testing at GRC under conditions simulating the structural and dynamic environment encountered during engine operation. AFRL (WPAFB) supported this testing by performing 3D laser vibrometry to identify the mixer mode shapes and modal frequencies. The successful fabrication and testing of such a component has been achieved. The CMC mixer demonstrated good durability during vibration testing at room and elevated temperature (TRL5). This has cleared the article for a ground-based test on a Rolls-Royce AE3007 engine, where the performance and benefits of the component can be further assessed.

exhaust nozzles↗