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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Interference Drag Associated with Engine Locations for Multidisciplinary Design Optimization

This research aims to quantify the interference drag for various engine locations on a traditional “tube and wing,” 150-passenger commercial aircraft flying at 35,000 ft and Mach 0.8. Engine locations are varied in the chord wise, span wise, and vertical directions near the wing, both under and above the wing, as well as along the fuselage. Euler simulations are performed with representative powered modern engines. The results are intended to supplement empirical drag estimates suitable for multidisciplinary design environments. Large interference drag increases, as compared to the isolated air frame and engine geometry, are found to occur when the engine is placed directly above or below the wing. Interference effects are significantly reduced, and in some instances result in benefits compared to the isolated bodies, when the engines are placed fore or aft of the wing. Interference drag increases are partially explained by flow channels leading to choked flow and shock interactions between bodies.

Blaesser, Nathaniel J.↗
NASA NTRS1982

Flows through sequential orifices with heated spacer reservoirs

Flow rates and pressure thermal profiles for two phase choked flows of fluid nitrogen were studied theoretically and experimentally in a four sequential orifice configuration. Both theory and experimental evidence demonstrate that heat addition in the first spacer-reservoir adjacent to the inlet orifice is most effective in reducing the flow rate and that heat addition in the last spacer-reservoir is least effective. The flows are choked at the exit orifice for large spacings and at the inlet orifice for small spacings. The moderate addition of heat available for this experiment did not materially alter this result for large spacings; however, significant heat addition for the small spacings tended to shift the choke point to the exit orifice. Nitrogen is used as the working fluid over a range of states from liquid to gas with a reduced inlet stagnation pressure range to P sub r, o = 2.

Hendricks, R. C.↗
NASA NTRS1972

Quasi-one-dimensional compressible flow across face seals and narrow slots. 2: Computer program

A computer program is presented for compressible fluid flow with friction across face seals and through narrow slots. The computer program carries out a quasi-one-dimensional flow analysis which is valid for laminar and turbulent flows under both subsonic and choked flow conditions for parallel surfaces. The program is written in FORTRAN IV. The input and output variables are in either the International System of Units (SI) or the U.S. customary system.

Zuk, J.↗
NASA NTRS1975

Compressible flow computer program for gas film seals

Computer program, AREAX, calculates properties of compressible fluid flow with friction and area change. Program carries out quasi-one-dimensional flow analysis which is valid for laminar and turbulent flows under both subsonic and choked flow conditions. Program was written to be applied to gas film seals.

Zuk, J.↗
NASA NTRS1983

Analogy between fluid cavitation and fracture mechanics

When the stresses imposed on a fluid are sufficiently large, rupture or cavitation can occur. Such conditions can exist in many two-phase flow applications, such as the choked flows, which can occur in seals and bearings. Nonspherical bubbles with large aspect ratios have been observed in fluids under rapid acceleration and high shear fields. These bubbles are geometrically similar to fracture surface patterns (Griffith crack model) existing in solids. Analogies between crack growth in solid and fluid cavitation are proposed and supported by analysis and observation (photographs). Healing phenomena (void condensation), well accepted in fluid mechanics, have been observed in some polymers and hypothesized in solid mechanics. By drawing on the strengths of the theories of solid mechanics and cavitation, a more complete unified theory can be developed.

Hendricks, R. C.↗
NASA NTRS1950

Investigation of Performance of Axial-Flow Compressor of XT-46 Turbine-Propeller Engine. I - Preliminary Investigation at 50-,70-, and 100-Percent Design Equivalent Speed

An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations. The station near the entrance section of the combustors indicated a peak pressure ratio of 6.3 at an adiabatic efficiency of 0.63 for a corrected weight flow of 23.1 pounds per second. The other, located one blade-chord downstream of the last stator row, indicated a peak pressure ratio of 6.97 at an adiabatic efficiency of 0.81 for a corrected weight flow of 30.4 pounds per second. The difference in performance obtained at the two stations is attributed to shock waves in the vicinity of the last stator row. These shock waves and the accompanying flow choking, together with interstage circulatory flows, shift the compressor operating curves into the region where surge would normally occur. The inability of the compressor to meet design pressure ratio is probably due to boundary-layer buildup in the last stages, which cause axial velocities greater than design values that, in turn, adversely affect the angles of attack and turning angles in these blade rows.

John W R Creagh↗
NASA NTRS2003

High-Flow Jet Exit Rig Designed and Fabricated

The High-Flow Jet Exit Rig at the NASA Glenn Research Center is designed to test single flow jet nozzles and to measure the appropriate thrust and noise levels. The rig has been designed for the maximum hot condition of 16 lbm/sec of combustion air at 1960 R (maximum) and to produce a maximum thrust of 2000 lb. It was designed for cold flow of 29.1 lbm/sec of air at 530 R. In addition, it can test dual-flow nozzles (nozzles with bypass flow in addition to core flow) with independent control of each flow. The High- Flow Jet Exit Rig was successfully fabricated in late 2001 and is being readied for checkout tests. The rig will be installed in Glenn's Aeroacoustic Propulsion Laboratory. The High-Flow Jet Exit Rig consists of the following major components: a single component force balance, the natural-gas-fueled J-79 combustor assembly, the plenum and manifold assembly, an acoustic/instrumentation/seeding (A/I/S) section, a table, and the research nozzles. The rig will be unique in that it is designed to operate uncooled. The structure survives the 1960 R test condition because it uses carefully selected high temperature alloy materials such as Hastelloy-X. The lower plenum assembly was designed to operate at pressures to 450 psig at 1960 R, in accordance with the ASME B31.3 piping code. The natural gas-fueled combustor fires directly into the lower manifold. The hot air is directed through eight 1-1/2-in. supply pipes that supply the upper plenum. The flow is conditioned in the upper plenum prior to flowing to the research nozzle. The 1-1/2-in. supply lines are arranged in a U-shaped design to provide for a flexible piping system. The combustor assembly checkout was successfully conducted in Glenn's Engine Component Research Laboratory in the spring of 2001. The combustor is a low-smoke version of the J79 combustor used to power the F4 Phantom military aircraft. The natural gas-fueled combustor demonstrated high-efficiency combustion over a wide range of operating conditions. This wide operating envelope is required to support the testing of both single- and dual-flow nozzles. Key research goals include providing simultaneous, highly accurate acoustic, flow, and thrust measurements on jet nozzle models in realistic flight conditions, as well as providing scaleable acoustic results. The High-Flow Jet Exit Rig is a second-generation high-flow test rig. Improvements include cleaner flow with reduced levels of particulate, soot, and odor. Choked-flow metering is required with plus or minus 0.25-percent accuracy. Thrust measurements from 0 to 2000 lbf are required with plus or minus 0.25-percent accuracy. Improved acoustics will be achieved by minimizing noise through large pipe bend radii, lower internal flow velocities, and microdrilled choke plates with thousands of 0.040-in.- diameter holes.

Buehrle, Robert J.↗
NASA NTRS1981

Finite element analysis of transonic flows in highly staggered cascades

Finite element method is employed in the solution of full-potential equation for analyzing steady, transonic flows through a series of highly staggered cascades. Convergence characteristics and efficiency of the developed numerical procedure are discussed. Accuracy of the mathematical model is investigated in terms of the employed computational grid. In particular, the importance of accurate representation of the trailing edge together with the shock on the obtained flow results are studied. Numerical results include choked flows, for a cascade of NACA 0012 airfoils with pitch/chord length = 1, and stagger angle of 45 deg.

Akay, H. U.↗
NASA NTRS1983

End-wall boundary layer measurements in a two-stage fan

Detailed flow measurements made in the casing boundary layer of a two-stage transonic fan are summarized. These measurements were taken at stations upstream of the fan, between all blade rows, and downstream of the last blade row. At the design tip speed of 429 m/sec the fan achieved a peak efficiency of 0.846 at a pressure ratio of 2.471. The boundary layer data were obtained at three weight flows at the design speed: one near choke flow, one near peak efficiency, and one near stall. Conventional boundary layer parameters were calculated from the data measured at each measuring station for each of the three flows. A classical two dimensional casing boundary layer was measured at the fan inlet and extended inward to approximately 15 percent of span. A highly three dimensional boundary layer was measured at the exit of each blade row and extended inward to approximately 10 percent of span. The steep radial gradient of axial velocity noted at the exit of the rotors was reduced substantially as the flow passed through the stators. This reduced gradient is attributed to flow mixing. The amount of flow mixing was reflected in the radial redistribution of total temperature as the flow passed through the stators. The data also show overturning of the tip flow at the stator exits that is consistent with the expected effect of the secondary flow field. The blockage factors calculated from the measured data show an increase in blockage across the rotors and a decrease across the stators.

Ball, C. L.↗
NASA NTRS2004

Performance of Thermal Mass Flow Meters in a Variable Gravitational Environment

The performance of five thermal mass flow meters, MKS Instruments 179A and 258C, Unit Instruments UFM-8100, Sierra Instruments 830L, and Hastings Instruments HFM-200, were tested on the KC-135 Reduced Gravity Aircraft in orthogonal, coparallel, and counterparallel orientations relative to gravity. Data was taken throughout the parabolic trajectory where the g-level varied from 0.01 to 1.8 times normal gravity. Each meter was calibrated in normal gravity in the orthogonal position prior to flight followed by ground testing at seven different flow conditions to establish a baseline operation. During the tests, the actual flow rate was measured independently using choked-flow orifices. Gravitational acceleration and attitude had a unique effect on the performance of each meter. All meters operated within acceptable limits at all gravity levels in the calibrated orthogonal position. However, when operated in other orientations, the deviations from the reference flow became substantial for several of the flow meters. Data analysis indicated that the greatest source of error was the effect of orientation, followed by the gravity level. This work emphasized that when operating thermal flow meters in a variable gravity environment, it is critical to orient the meter in the same direction relative to gravity in which it was calibrated. Unfortunately, there was no test in normal gravity that could predict the performance of a meter in reduced gravity. When operating in reduced gravity, all meters indicated within 5 percent of the full scale reading at all flow conditions and orientations.

Brooker, John E.↗
NASA NTRS2017

Low Pressure Flame Blowoff from the Forward Stagnation Region of a Blunt-Nosed Cast PMMA Cylinder in Axial Mixed Convective Flow

Low-pressure blowoff experiments were conducted with a stagnation flame stabilized on the forward tip of cast PMMA rods in a vertical wind tunnel. Pressure, forced flow velocity, gravity, and ambient oxygen concentration were varied. Stagnation flame blowoff is determined from a time-stamped video recording of the test. The blowoff pressure is determined from test section pressure transducer data that is synchronized with the time stamp. The forced flow velocity is also determined from the choked flow orifice pressure. Most of the tests were performed in normal gravity, but a handful of microgravity tests were also conducted to determine the influence of buoyant flow velocity on the blowoff limits. The blowoff limits are found to have a linear dependence between the partial pressure of oxygen and the total pressure, regardless of forced flow velocity and gravity level. The flow velocity (forced and/or buoyant) affects the blowoff pressure through the critical Damkohler number residence time, which dictates the partial pressure of oxygen at blowoff. This is because the critical stretch rate increases linearly with increasing pressure at low pressure (sub-atmospheric pressures) since a second-order overall reaction rate with two-body reactions dominates in this pressure range.

flammability↗
NASA NTRS2016

Unsteady Flow Simulation of a Sweeping Jet Actuator Using a Lattice-Boltzmann Method

Active flow control technology is increasingly used in aerospace applications to control flow separation and to improve aerodynamic performance. In this paper, PowerFLOW is used to simulate the flow through a sweeping jet actuator at two different pressure ratios. The lower pressure ratio leads to a high subsonic flow, whereas the high pressure ratio produces a choked flow condition. Comparison of numerical results with experimental data is shown, which includes qualitatively good agreement of pressure histories and spectra. PIV measurements are also available but the simulation overestimates mean and fluctuation quantities outside the actuator. If supply pressure is matched at one point inside the mixing chamber a good qualitative agreement is achieved at all other monitor points.

Duda, B.↗
NASA NTRS2023

Benchmark Performance Metrics of a Vane Propellant Management Device for a 0.15 m3 Liquid Hydrogen Tank

The use of cryogenic propellants has and will continue to play an integral role in manned-space exploration due the high specific impulses offered and its ubiquity through in-situ resource utilization. But guaranteeing vapor-free transfer of such low-surface tension liquids is difficult for traditional capillary-action propellant management devices (PMDs). Screen-channel liquid acquisition devices and compliant origami bladders are potential solutions, but to quantify the benefits these technologies offer, this paper present a cast study analyzing the performance metrics of an orthodox vane PMD as a benchmark for comparison. A 0.15 m3 liquid hydrogen tank at 20.3 K and 103 kPa was selected for study. Then assuming no body forces and no heat transfer (for simplicity), a steady-state 1-D differential equation was numerically solved in tandem with four possible wetted area configurations to yield an expulsion efficiency for a theoretically maximum inputted expulsion flow rate. Maximization of the flow rate was constrained by the onset of choked flow. Additional inputs, including vane height and vane number, were parametrically varied between 0.1 – 10 cm and 4 – 28 vanes, respectively.

Liquid Acquisition Device↗
NASA NTRS1981

Flow through axially aligned sequential apertures of the orifice and Borda types

Choked flow rate and pressure profile data were obtained and studied for two axially aligned sequential configurations consisting of four Borda type inlets of 1.9 1/D with two separation distances of 0.8 and 30 diameters and four orifice type inlets of 0.5 1/D with two separation distances of 0.66 and 32 diameters. Data were obtained using fluid nitrogen over the reduced inlet temperature and pressure range 0.68 T/T sub c gas and P/P sub c to 2. A flow coefficient reduced temperature plot can be used to represent the flow rate data for each geometry. At the larger separation distances, the pressure profiles dropped sharply at the entrance and partially recovered within each of the Borda and orifice inlet configurations; the exception being the last inlet where at low entrance temperatures, fluid jetting could occur. For the smaller spacings fluid jetting was prevalent throughout each of the inlet configurations at lower inlet temperatures. These results are in qualitative agreement with data of tubes with single Borda or sharp edge orifice type inlets to 105 1/D and water flow visualization studies.

Hendricks, R. C.↗
NASA NTRS1981

Some flow phenomena associated with aligned, sequential apertures with Borda-type inlets

Choked flow rate and pressure profile data were taken and studied for a configuration consisting of four axially aligned, sequential Borda tubes of 1.9 length diameter ratio with separation distances of 0.8 and 30 tube diameters. For either case the flow rate data could be represented by a flow coefficient reduced temperature plot. At a separation distance of 30 tube diameters the pressure profiles dropped sharply at the entrance and recovered within each Borda tube; except at low temperatures, where fluid jetting through the last Borda tube occurred. At a separation distance of 0.8 tube jetting was prevalent, and application of a significant backpressure did not alter the jetting. These results agree with other data for tubes with Borda or sharp edge orifice inlets and with a water flow visualization study reported herein.

Hendericks, R. C.↗
NASA NTRS1981

Experiments on flow through one to four inlets of the orifice and Borda type

Choked flow rate and pressure profile data were taken on sequential axially aligned inlets of the orifice and Borda type. The configuration consisted of from two to four inlets spaced at two nominal separation distances of 0.7 and 30 diameters. At the nominal 30 diameter spacing, the reduced flow rate follows a simple empirical relation based on the reduced flow rate for a single inlet. At the nominal 0.7 diameter spacing, fluid jetting was prevalent at low temperatures and flow rates were the same as for a single inlet.

Hendricks, R. C.↗
NASA NTRS1981

Flow through axially aligned sequential apertures of the orifice and Borda types

Choked flow rate and pressure profile data were taken and studied for two axially aligned sequential configurations consisting of: (1) Four Borda type inlets of 1.9 1/D with two separation distances of 0.8 and 30 diameters. (2) Four orifice type inlets of 0.5 1/D with two separation distances of 0.66 and 32 diameters. A flow-coefficient reduced-temperature plot can be used to represent the flow rate data for each geometry. At the larger separation distances, the pressure profiles dropped sharply at the entrance and partially recovered within each of the Borda and orifice inlet configurations; the exception being the last inlet where at low entrance temperatures, fluid jetting could occur. For the smaller spacings fluid jetting was prevalent throughout each of the inlet configurations at lower inlet temperatures. These results are in qualitative agreement with data of tubes with single Borda or sharp-edge orifice type inlets to 105 1/D and water flow visualization studies.

Hendricks, R. C.↗