Search NASA⌕ Search

SEARCH · Search NASA

Results for “Absolute pressure”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

An occultation satellite system for determining pressure levels in the atmosphere

A two-satellite microwave occultation system is described that will fix, as an absolute function of altitude, the pressure-temperature profile generated by a passive infrared sounder. The 300 mb pressure level is determined to within 24 m rms, assuming the temperture errors produced by the infrared sensor are not greater than 2 K rms. Error caused by water vapor in the radio path is corrected by climatological adjustments. A ground test of the proposed system is described. A microwave signal propagating between two mountain tops was found to be subject to periods of intense fading. Computer analysis of the raypath between the transmitting and receiving stations indicates that multipath and defocusing were responsible for this fading. It is unlikely that an operational pressure-reference-level system will be subject to the deep fades observed in the ground test, because the phenomena are associated with lower altitudes than the closest approach altitude of an occultation-system raypath.

Morrison, A. R.↗
NASA NTRS1948

Performance of J-33-A-21 Turbojet-Engine Compressor I - Over-All Performance Characteristics at Equivalent Impeller Speeds from 6000 to 13,400 RPM

The NACA is investigating a series of J-33 turbojet-engine compressors to determine the over-all and component performances and to improve theories of flow through large centrifugal compressors, The production model J-33-A-21 was operated over a range of inlet temperatures from 80 to -40 F and inlet pressures from 14 to 5 inches mercury absolute for equivalent impeller speeds from 6000 to 13,400 rpm. At the equivalent design speed of 11,500 rpm, the compressor had a peak pressure ratio of 3.98 at an equivalent weight flow of 73.4 pounds per second and an adiabatic temperature-rise , efficiency of 0.701. When the compressor speed was reduced from the design speed to 6000 rpm, the adiabatic temperature-rise efficiency increased to 0.747. At the maximum equivalent speed investigated (13,400 rpm), a peak pressure ratio of 5.09 was obtained at an adiabatic temperature-rise efficiency of 0.617 and an equivalent weight flow of 66.O pounds per second. An increase in inlet pressure from 5.5 to 14 inches mercury absolute, with a consequent increase in Reynolds number index, improved the pressure ratio but had no apparent effect on the ratio of temperature rise through the compressor to inlet temperature. The variation of the peak adiabatic temperature-rise efficiency with inlet pressure is in the direction that would be expected from a Reynolds number effect. Decrease in the inlet temperature from 80 to -40 F, with a consequent increase in Reynolds number index, resulted in scatter of the pressure-ratio data and increased values of temperature ratio. The variation of the adiabatic temperature-rise efficiency with inlet temperature is probably the result of heat-transfer effects and scatter in the pressure ratio.

Beede, William L.↗
DOE OSTI2021

Impact of low reactivity fuel type on low load combustion, emissions, and cyclic variations of diesel-ignited dual fuel combustion

In this study, cyclic variations in dual fuel combustion with diesel ignition of three different low reactivity fuels (methane, propane, and gasoline) are examined under identical operating conditions. Experiments were performed on a single cylinder research engine (SCRE) at a low load of 3.3 bar brake mean effective pressure (BMEP). The start of injection (SOI) of diesel was varied from 280 to 330 absolute crank angle degrees (CAD). Engine speed, rail pressure, and boost pressure were held constant at 1500 rpm, 500 bar, and 1.5 bar, respectively. The energy substituted by the low reactivity fuel was fixed at 80% of the total energy input. It was found that diesel-methane (DM) and diesel-propane (DP) combustion were affected by diesel mixing to a greater extent than diesel-gasoline (DG) combustion due to the higher reactivity of gasoline. The magnitude of low temperature heat release was greatest for DG combustion followed by DM and DP combustion for all SOIs. The ignition delay for DG combustion was the shortest, followed by DM and DP combustion. DM and DP combustion exhibited more cyclic variations than DG combustion. Cyclic variations decreased for DM and DP combustion when SOI was advanced; however, DG combustion cyclic variations remained essentially constant for all SOIs. Earlier SOIs (280, 290, 300, and 310 CAD) for DM and (280, 290, and 300 CAD) for DP combustion indicated some prior-cycle effects on the combustion and IMEP (i.e. some level of determinism).

Engineering↗
NASA NTRS1960

Performance of Several Method-of-Characteristics Exhaust Nozzles

Nozzle performance data were obtained with three "method-of-characteristics" nozzles and a 150 conical nozzle at pressure ratios up to 130. Each basic configuration was cut off and tested at expansion ratios of 25, 20, 15, and 10. Unheated dry air was used at nozzle inlet pressures up to 22,000 pounds per square foot absolute. Nozzle thrust data were extrapolated to infinite pressure ratio (zero discharge pressure). As much as 1-percent increase in thrust with no increase in nozzle surface area (weight), can be obtained by using a method-of-characteristics, nozzle instead of a 15 conical nozzle when operating with a nozzle expansion ratio of 25 and nozzle pressure ratios from 200 to infinity. Conversely, for the same thrust, reductions in nozzle divergent surface area in the order of 25 percent are possible. The thrust performance of the method-of-characteristics nozzle was not as good as that of the 150 conical nozzle when operating at pressure ratios considerably below design (below 100 for the expansion ratio 25 nozzles). Theoretical and measured nozzle momentum coefficients agreed within about 0.6 percent. This is the order of accuracy of both the measured and theoretical values.

Farley, John M.↗
NASA NTRS2008

Prebreathe Protocol for Extravehicular Activity Technical Consultation Report

In the performance of EVA by that National Aeronautics and Space Administration (NASA) astronauts, there exists a risk of DCS as the suit pressure is reduced to 4.3 pounds per square inch, absolute (psia) from the International Space Station (ISS) pressure of 14.7 psia. Several DCS-preventive procedures have been developed and implemented. Each of these procedures involve the use of oxygen (O2) prebreathe to effectively washout tissue nitrogen (N2).The management of the ISS Programs convened an expert independent peer review Team to conduct a review of the Decompression Sickness (DCS) risks associated with the Extra Vehicular Activity (EVA) Campout Prebreathe (PB) protocol for its consideration for use on future missions. The major findings and recommendations of the expert panel are: There is no direct experimental data to confirm the potential DCS risks of the Campout PB protocol. However, based on model data, statistical probability, physiology, and information derived from similar PB protocols, there is no compelling evidence to suggest that the Campout PB protocol is less safe than the other NASA approved PB protocols.

Ross, Jerry↗
NASA NTRS2009

Aerojet - AJ10-137 Apollo Service Module Engine

The general configuration of the SPS engine was 20,000 pounds of thrust, with a chamber pressure of 100 psi and specific impulse (Isp) of 314.5. The very large nozzle had an area ratio of 62.5:1 (exit area to throat area). The propellants were nitrogen tetroxide (also known as N2O4 and nitrous oxide) and A-50. A-50 was a hydrazine family fuel. Aerojet developed it for the Titan Missile Program when they went with Titan II, to store it in the launch silos. They wanted the highest performance they could get. N2H4 was just pure hydrazine, which doesn't take low temperature very well. In fact, it freezes about like water. We started adding unsymmetrical-dimethylhydrazine (UDMH) to the hydrazine until such time as it would meet the environmental specifications the Air Force needed for Titan II. It turned out it s roughly a fifty-fifty mix. We still had to be careful with that fuel because the two fluids didn't mix very well chemically. We had to spray the two fluids through some special nozzles to get them to emulsify with each other into a single fluid. If we ever got it too cold or froze it, the hydrazine separated back out. Then, if we tried to run the engine, things could go boom in the night. The inlet pressure was only 165 pounds per square inch absolute (psia), but we needed at least forty psi pressure drop across the injector just to get some kind of stable flow. It was a whole new game for some of us. We didn't have much supply pressure to work with. It had the aluminum injector to keep the weight down. That was a couple feet in diameter, and we didn't have a lot of propellant to cool it. In fact, we had to use both propellants to keep the injector cool. There were twenty-two ring channels in the injector. Specification required 750 seconds duration, or fifty engine restarts during a flight. There were several first flight things we accomplished with the engine. It was the first ablative thrust chamber of any size to fly. (See Slide 6, Appendix G) There were no liners in it. It was just straight ablative material. It took us a while to figure that out. It was a throat-gimbaled engine, and it was the first engine to fly with columbium (also known as niobium, used as an alloying element in steels and superalloys) in the nozzle.

Boyce, Clay↗
NASA NTRS1948

Performance of Allison Model 400-C6 Turbojet-Engine Compressor: Over-All Performance Characteristics of Compressor at Equivalent Impeller Speeds of 6000, 7000, and 8500 RPM - I

The Allison model 400-C6 compressor was operated at an inlet pressure of 12 inches of mercury absolute ana ambient inlet temperature at equivalent impeller speeds of 6000, 7000, and 8500 rpm. Additional runs at an equivalent speed of 7000 rpm and ambient inlet temperature were made at inlet pressures from 7 to 22 inches of mercury absolute. The results of this investigation are compared with those of the 533-A-23 compressors. For the speeds investigated, the Allison model 400-C6 compressor had a maximum adiabatic temperature-rise efficiency of 0.768 at an equivalent speed of 7000 rpm; the corresponding equivalent weight flow was 45.0 pounds per second and the pressure ratio was 1.83. At an equivalent impeller speed of 8500 rpm, the maximum equivalent weight flow was 61.6 pounds per second and the peak pressure ratio of 2.38 occurred at an equivalent weight flow of 52.2 pounds per 1 second and an adiabatic temperature-rise efficiency of 0.714. At an equivalent speed of 7000 rpm, increasing the compressor- inlet pressure increased the maximum equivalent weight flow and the pressure ratio.

Kovach, Karl↗
NASA NTRS1961

Annular Internal-External-Expansion Rocket Nozzles for Large Booster Applications

For large-thrust booster applications, annular rocket nozzles employing both internal and external expansion are investigated. In these nozzles, free-stream air flows through the center as well as around the outside of the exiting jet. Flaps for deflecting the rocket exhaust are incorporated on the external-expansion surface for thrust-vector control. In order to define nozzle off-design performance, thrust vectoring effectiveness, and external stream effects, an experimental investigation was conducted on two annular nozzles with area ratios of 15 and 25 at Mach 0, 2, and 3 in the Lewis 10- by 10-foot wind tunnel. Air, pressurized to 600 pounds per square inch absolute, was used to simulate the exhaust flow. For a nozzle-pressure-ratio range of 40 to 1000, the ratio of actual to ideal thrust was essentially constant at 0.98 for both nozzles. Compared with conventional convergent-divergent configurations on hypothetical boost missions, the performance gains of the annular nozzle could yield significant orbital payload increases (possibly 8 to 17 percent). A single flap on the external-expansion surface of the area-ratio-25 annular nozzle produced a side force equal to 4 percent of the axial force with no measurable loss in axial thrust.

Connors, James F.↗
NASA NTRS1959

Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation

A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.

Sivo, Joseph N.↗
NASA NTRS1985

Middle Atmosphere Reference Model Derived from Satellite Data

Zonally averaged temperature and geostrophic zonal wind for each month and for latitudes from 80 S to 80 N with pressure scale height as a vertical coordinate are given tabular form. The pressure scale height is defined as -ln (p/P sub O) where p is pressure and P sub O is surface pressure. The height interval corresponding to one pressure scale height is proportional to absolute temperature, and is 7 km at 240 K. Values in are given at intervals of 0.5 pressure scale heights, i.e., approximately 3,5 km. The geopotential height fields for the principal seasons are given. Temperature, pressure and density with geometric height as the vertical coordinate at invervals of 5 km. The temperature values were obtained from a combination of satellite data above 30 mb with values supplied by the Berlin Free University at 30 mb and the climatology derived by OORT (1983) for 50 mb and below. The geopotential height fields were obtained from these temperature fields by integrating up and down from the 30 mb geopotential height supplied by Berlin Free University. The geostrophic winds were obtained by differentiating these geopotential height fields.

Barnett, J. J.↗
NASA NTRS1987

Measuring Viscosities of Gases at Atmospheric Pressure

Variant of general capillary method for measuring viscosities of unknown gases based on use of thermal mass-flowmeter section for direct measurement of pressure drops. In technique, flowmeter serves dual role, providing data for determining volume flow rates and serving as well-characterized capillary-tube section for measurement of differential pressures across it. New method simple, sensitive, and adaptable for absolute or relative viscosity measurements of low-pressure gases. Suited for very complex hydrocarbon mixtures where limitations of classical theory and compositional errors make theoretical calculations less reliable.

Singh, Jag J.↗
NASA NTRS1938

Engine Performance and Knock Rating of Fuels for High-output Aircraft Engines

Data are presented to show the effects of inlet-air pressure, inlet-air temperature, and compression ratio on the maximum permissible performance obtained on a single-cylinder test engine with aircraft-engine fuels varying from a fuel of 87 octane number to one 100 octane number plus 1 ml of tetraethyl lead per gallon. The data were obtained on a 5-inch by 5.75-inch liquid-cooled engine operating at 2,500 r.p.m. The compression ratio was varied from 6.50 to 8.75. The inlet-air temperature was varied from 120 to 280 F. and the inlet-air pressure from 30 inches of mercury absolute to the highest permissible. The limiting factors for the increase in compression ratio and in inlet-air pressure was the occurrence of either audible or incipient knock. The data are correlated to show that, for any one fuel,there is a definite relationship between the limiting conditions of inlet-air temperature and density at any compression ratio. This relationship is dependent on the combustion-gas temperature and density relationship that causes knock. The report presents a suggested method of rating aircraft-engine fuels based on this relationship. It is concluded that aircraft-engine fuels cannot be satisfactorily rated by any single factor, such as octane number, highest useful compression ratio, or allowable boost pressure. The fuels should be rated by a curve that expresses the limitations of the fuel over a variety of engine conditions.

Rothbrock, A M↗
DOE OSTI2022

Experimental Studies of Low-Load Limit in a Stoichiometric Micro-Pilot Diesel Natural Gas Engine

While operating at light loads, diesel pilot-ignited natural gas engines with lean premixed natural gas suffer from poor combustion efficiency and high methane emissions. This work investigates the limits of low-load operation for a micro-pilot diesel natural gas engine that uses a stoichiometric mixture to enable methane and nitrogen oxide emission control. By optimizing engine hardware, operating conditions, and injection strategies, this study focused on defining the lowest achievable load while maintaining a stoichiometric equivalence ratio and with acceptable combustion stability. A multi-cylinder diesel 6.7 L engine was converted to run natural gas premix with a maximum diesel micro-pilot contribution of 10%. With a base diesel compression ratio of 17.3:1, the intake manifold pressure limit was 80 kPa(absolute). At a reduced compression ratio of 15:1, this limit increased to 85 kPa, raising the minimum stable load. Retarding the combustion phasing, typically used in spark-ignition engines to achieve lower loads, was also tested but found to be limited by degraded diesel ignition at later timings. Reducing the pilot injection pressure improved combustion stability, as did increasing pilot quantity at the cost of lower substitution ratios. The lean operation further reduced load but increased NOx and hydrocarbon emissions. At loads below the practical dual-fuel limit, a transition to lean diesel operation will likely be required with corresponding implications for the aftertreatment system.

03 NATURAL GAS↗
DOE OSTI2024

A miniature jet pump for slurries

Here, this work reports on the design and performance of a jet pump system to pump slurries. The jet pump uses a high-velocity water jet to entrain and pump the slurry. This pumping action requires no valves and does not use moving solid parts. Hence, the system can be readily sealed, and the pumping action creates no observable mechanical damage to slurry particles. We tested carbon black beads suspended in water at concentrations of 0.0, 0.2, and 2.0 wt% at primary jet flow rates of 50 and 100 mL/min. Our measurements show an approximately linear decrease in entrainment ratio (slurry-to-jet flow rate) and absolute slurry flow rate with increasing adverse (load) pressure difference. The entrainment ratio is a strong function of the absolute jet flow rate. Pump performance was insensitive to slurry composition. Micron-resolution particle image velocimetry (micro-PIV) revealed qualitative changes in the pump’s internal flow due to changes in absolute jet flow rate, including evidence of the suppression of a vortex at higher flow rates. These findings establish key operational parameters for the jet pump and highlight its potential in a variety of slurry flow pumping applications. The pump is particularly suited for flow electrode slurries (e.g. as used for energy storage and water treatment) as it maintains particle integrity and can be used with standard tubing, tube fittings, and sealing methods.

42 ENGINEERING↗
NASA NTRS1946

Investigation of Operating Characteristics of an Engine Equipped with Modifications to Eliminate Fuel-Evaporation Icing

Two modified fuel-injection systems, a drilled-inducer type and a spinner type, that prevent serious fuel-evaporation icing were installed on a V-type, liquid-cooled aircraft engine and a preliminary investigation was conducted to determine the effect on engine operating characteristics. The spinner system was also ground - and flight tested on a twin-engine fighter airplane. Flight measurements of cylinder-head temperature over a range of fuel-air ratios and engine power conditions were made at an altitude of approximately 10,000 feet. Starting and accelerating of the engine on the ground were unaffected by the fuel-injection modifications. During the flight investigation, no appreciable variation occurred between the maximum and minimum cylinder-head temperature with the standard and modified system for the same power condition and no irregularity of mixture distribution could be detected throughout the power range of the engine. Normal mixture distribution was also indicated by a similar response of cylinder-head temperature for variations of fuel-air ratio at manifold pressures of 25 and 35 inches of mercury absolute. Both modified fuel-injection systems required less fuel-nozzle pressure than the standard system to obtain the desired fuel-air ratio for given air-flow condition.

Mulholland, Donald R.↗
DOE OSTI2025

CHELAX-BNF: Fast Response Temperature by airborne measurements

The original data were collected on board the ARM Aerial Facility ArcticShark uncrewed aerial system (UAS; https://www.arm.gov/capabilities/observatories/aaf/uas ) during the “Characterizing HEterogeneous Land-Atmosphere eXchanges at BNF” field campaign (CHEAX-BNF; https://arm.gov/research/campaigns/aaf2025CHELAX-BNF ). The ARM Aerial Facility ArcticShark UAS was based at the public-use airport of Posey Field, AL (FAA LID: 1M4, 34.28027778° N, 87.60055556° W, 283m MSL) from May 28 to June 23, 2025. The ArcticShark UAS performed 5 flights, including 4 research flights over the BNF Main Site (ARM Mobile Facility 3, https://arm.gov/capabilities/observatories/amf ) and Supplemental Facilities to measure atmospheric state, turbulence, surface IR temperature and imagery, aerosol number concentration, and aerosol size distribution. The current data set presents fast response temperature in the atmospheric boundary layer and lower free troposphere measured on the airborne platform throughout the field campaign. The primary instruments used to create the current data set were the fine wire thermocouple probe, the Aircraft Integrated Meteorological Measurement System (AIMMS-30), the Pitot-static system (part of UAS flight control), and the infrared gas analyzer sensor for H2O and CO2 (LI-840). All parameters used in temperature calculations (static pressure, True Air Speed, and absolute humidity in the form of dew point temperature) were included in the data set. For user convenience, one additional parameter was also included: the type of flight flag (level, up, down, turn, and combination of thereof).

Air temperature, fast response↗