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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1948

Preliminary Results of Nene II Engine Altitude-chamber Performance Investigation: Altitude Performance using 18.00-inch-diameter Jet Nozzle - 3

An investigation is being conducted to determine the altitude performance characteristics of the Nene II engine and its components. The present paper presents the preliminary results obtained using jet nozzle 18.00 inches in diameter, with an area equal to 92.2 percent of the area of the standard jet nozzle for this engine. The experimental results presented are for conditions simulating altitudes from 20,000 to 60,000 feet and ram-pressure ratios from 1.1 to 3.5. These ram-pressure ratios correspond to flight Mach numbers between 0.374 and 1.466. Data obtained with the 18.00 inch-diameter jet nozzle and corrected to standard sea-level conditions showed substantially the same trends with altitude as the data previously obtained with an 18.75-inch-diameter nozzle and with an 18.41-inch-diameter nozzle. Jet thrust, air consumption, and fuel consumption, corrected to standard sea-level conditions, increased rapidly with increasing ram-pressure ratio. In general, corrected net thrust specific fuel consumption increased with increase in ram-pressure ratio. Corrected net thrust decreased with an increase in ram-pressure ratio at an engine speed of 8000 rpm. At corrected engine speeds between 8000 and 10,800 rpm, net thrust first decreased with an increase in ram-pressure ratio and then increased with further increase in ram pressure ratio; at corrected engine speeds above 10,800 rpm, net thrust increased continuously with increase in ram-pressure ratio. Tail-pipe temperature decreased with an increase in ram-pressure ratio.

Grey, Ralph E.↗
NASA NTRS2012

OH PLIF Visualization of the UVa Supersonic Combustion Experiment: Configuration A

Hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were performed in the University of Virginia s dual-mode scramjet experiment. The test section was set up in configuration A, which includes a Mach 2 nozzle, combustor, and extender section. Hydrogen fuel was injected through an unswept compression ramp at two different equivalence ratios. Through the translation of the optical system and the use of two separate camera views, the entire optical range of the combustor was accessed. Single-shot, average, and standard deviation images of the OH PLIF signal are presented at several streamwise locations. The results show the development of a highly turbulent flame structure and provide an experimental database to be used for numerical model assessment.

Johansen, Craig T.↗
NASA NTRS2013

OH PLIF Visualization of the UVa Supersonic Combustion Experiment: Configuration A

Hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were performed in the University of Virginia supersonic combustion experiment. The test section was set up in configuration A, which includes a Mach 2 nozzle, combustor, and extender section. Hydrogen fuel was injected through an unswept compression ramp at two different equivalence ratios. Through the translation of the optical system and the use of two separate camera views, the entire optically accessible range of the combustor was imaged. Single-shot, average, and standard deviation images of the OH PLIF signal are presented at several streamwise locations. The results show the development of a highly turbulent flame structure and provide an experimental database to be used for numerical model assessment.

Johansen, Craig T.↗
NASA NTRS2006

Fuel-Air Mixing and Combustion in Scramjets

At flight speeds, the residence time for atmospheric air ingested into a scramjet inlet and exiting from the engine nozzle is on the order of a millisecond. Therefore, fuel injected into the air must efficiently mix within tens of microseconds and react to release its energy in the combustor. The overall combustion process should be mixing controlled to provide a stable operating environment; in reality, however, combustion in the upstream portion of the combustor, particularly at higher Mach numbers, is kinetically controlled where ignition delay times are on the same order as the fluid scale. Both mixing and combustion time scales must be considered in a detailed study of mixing and reaction in a scramjet to understand the flow processes and to ultimately achieve a successful design. Although the geometric configuration of a scramjet is relatively simple compared to a turbomachinery design, the flow physics associated with the simultaneous injection of fuel from multiple injector configurations, and the mixing and combustion of that fuel downstream of the injectors is still quite complex. For this reason, many researchers have considered the more tractable problem of a spatially developing, primarily supersonic, chemically reacting mixing layer or jet that relaxes only the complexities introduced by engine geometry. All of the difficulties introduced by the fluid mechanics, combustion chemistry, and interactions between these phenomena can be retained in the reacting mixing layer, making it an ideal problem for the detailed study of supersonic reacting flow in a scramjet. With a good understanding of the physics of the scramjet internal flowfield, the designer can then return to the actual scramjet geometry with this knowledge and apply engineering design tools that more properly account for the complex physics. This approach will guide the discussion in the remainder of this section.

Drummond, J. Philip↗
NASA NTRS1987

Experimental evaluation of a translating nozzle sidewall radial turbine

Studies have shown that reduced specific fuel consumption of rotorcraft engines can be achieved with a variable capacity engine. A key component in such an engine in a high-work, high-temperature variable geometry gas generator turbine. An optimization study indicated that a radial turbine with a translating nozzle sidewall could produce high efficiency over a wide range of engine flows but substantiating data were not available. An experimental program with Teledyne CAE, Toledo, Ohio was undertaken to evaluate the moving sidewall concept. A variety of translating nozzle sidewall turbine configurations were evaluated. The effects of nozzle leakage and coolant flows were also investigated. Testing was done in warm air (121 C). The results of the contractual program were summarized.

Roelke, Richard J.↗
NASA NTRS1959

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Fortini, Anthony↗
DOE OSTI2023

Computational study of the ECN spray C via one-way coupling of internal nozzle flow and ensuing spray

Here, in this study, computational fluid dynamics simulations of the Engine Combustion Network Spray C injector were performed to investigate the effect of the internal nozzle flow on the development of the ensuing spray in automotive fuel injectors. A set of simulation best practices was first developed to achieve an accurate representation of the internal nozzle flow. Detailed, spatiotemporally defined information was extracted at the injector’s orifice exit and stored in maps that were subsequently employed to initialize a Lagrangian spray by means of a static, one-way coupling (OWC) approach. To highlight the impact of the in-nozzle flow on the spray, the OWC approach was compared with the conventional rate-of-injection (ROI) approach, for which the spray parcels are commonly initialized using a blob injection model. The results indicated that, using the ROI approach, considerable evaporation was observed within the plume in the near-nozzle region. On the other hand, the OWC approach showed evidence of evaporation farther downstream and at the plume periphery. The different behavior observed with the two approaches was mostly due to substantial differences in fuel mass distribution. The liquid fuel concentration predicted by the OWC approach matched well with X-ray tomography measurements of projected fuel mass per unit volume, in terms of both magnitude and distribution. In contrast, an unphysical distribution of fuel mass characterized by an extremely low liquid concentration was predicted by the ROI approach. In the far-field, the effect of the in-nozzle flow was found to be weaker as the spray penetrated and interacted with the ambient gas. In this region, the OWC and ROI approaches provided similar predictions of vapor penetration, spray morphology, and mixing-related characteristics.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1992

Nozzle material requirements and the status of intermetallic matrix composites

The HSCT exhaust nozzle must manage high temperature exhaust gases and pressure gradients while meeting HSCT economic and noise goals. The important features and requirements for an HSCT exhaust nozzle are shown for a 2DCD (two-dimensional convergent-divergent) design. The same requirements would apply to an axisymmetric design. Exhaust nozzle weight has an adverse effect on the overall aircraft range, payload, and engine specific fuel consumption and is therefore the primary driver for advanced exhaust nozzle materials. Because of the large airflow and pressure gradients, exhaust nozzles are extremely large and heavy when made from current materials. The use of advanced materials with higher specific strength will reduce the weight of exhaust nozzle components. In addition to the flow of high-temperature exhaust gases into the exhaust nozzle, ambient air is entrained to reduce gas exit velocities and suppress sound. This leads to components exposed to extremely high temperature gradients and, hence, high thermal stresses. Further, exhaust gases are highly oxidizing; material environmental resistance will be an important factor for long life. Several viable concepts have been identified to reduce noise through the mixture of exhaust and ambient air. Sound can be further suppressed by acoustic panels that absorb high-frequency noise.

Johnson, Andrew M.↗
NASA NTRS1931

Some Characteristics of Fuel Sprays at Low-injection Pressures

This report presents the results of tests conducted at the Langley Memorial Aeronautical Laboratory, Langley Field, Va., to determine some of the characteristics of the fuel sprays obtained from an 0.008-inch and a 0.020-inch open nozzle when injection pressures from 100 to 500 pounds per square inch were used. Fuel oil and gasoline were injected into air at densities of atmospheric land 0.325 pound per cubic foot. It was found that the penetration rate at these low pressures was about the same as the rate obtained with higher pressures. Spray cone-angles were small and individual oil drops were visible in all the sprays. Gasoline and fuel oil sprays had similar characteristics.

Rothrock, A M↗
NASA NTRS2014

Multi-Point Combustion System: Final Report

A low-NOx emission combustor concept has been developed for NASA's Environmentally Responsible Aircraft (ERA) program to meet N+2 emissions goals for a 70,000 lb thrust engine application. These goals include 75 percent reduction of LTO NOx from CAEP6 standards without increasing CO, UHC, or smoke from that of current state of the art. An additional key factor in this work is to improve lean combustion stability over that of previous work performed on similar technology in the early 2000s. The purpose of this paper is to present the final report for the NASA contract. This work included the design, analysis, and test of a multi-point combustion system. All design work was based on the results of Computational Fluid Dynamics modeling with the end results tested on a medium pressure combustion rig at the UC and a medium pressure combustion rig at GRC. The theories behind the designs, results of analysis, and experimental test data will be discussed in this report. The combustion system consists of five radially staged rows of injectors, where ten small scale injectors are used in place of a single traditional nozzle. Major accomplishments of the current work include the design of a Multipoint Lean Direct Injection (MLDI) array and associated air blast and pilot fuel injectors, which is expected to meet or exceed the goal of a 75 percent reduction in LTO NOx from CAEP6 standards. This design incorporates a reduced number of injectors over previous multipoint designs, simplified and lightweight components, and a very compact combustor section. Additional outcomes of the program are validation that the design of these combustion systems can be aided by the use of Computational Fluid Dynamics to predict and reduce emissions. Furthermore, the staging of fuel through the individually controlled radially staged injector rows successfully demonstrated improved low power operability as well as improvements in emissions over previous multipoint designs. Additional comparison between Jet- A fuel and a hydrotreated biofuel is made to determine viability of the technology for use with alternative fuels. Finally, the operability of the array and associated nozzles proved to be very stable without requiring additional active or passive control systems. A number of publications have been publish

Goeke, Jerry↗
NASA NTRS1995

Thrust measurement in a 2-D scramjet nozzle

The two-dimensional thrust nozzle presents a challenging problem. The loading is not axisymmetric as in the case of a cone and the internal flow presents some design difficulties. A two-sting system has been chosen to accomodate the internal flow and achieve some symmetry. The situation is complicated by the fact that with the small ramp angle and the internal pressure on the nozzle walls, loading is predominantly transverse. Yet it is the axial thrust which is to be measured (i.e., the tensile waves propagating in the stings). Although bending stress waves travel at most at only 60% of the speed of the axial stress waves, the system needs to be stiffened against bending. The second sting was originally only used to preserve symmetry. However, the pressures on each thrust surface may be quite different at some conditions, so at this stage the signals from both stings are being averaged as a first order approximation of the net thrust. The expected axial thrust from this nozzle is not large so thin stings are required. In addition, the contact area between nozzle and sting needs to be maximized. The result was that it was decided to twist the stings through 90 deg, without distorting their cross-sectional shape, just aft of the nozzle. Finite element analysis showed that this would not significantly alter the propagation of the axial stress wave in the sting, while the rigidity of the system is greatly increased. A Mach 4 contoured nozzle is used in the experiments. The thrust calculated by integrating the static pressure measurements on the thrust surfaces is compared with the deconvolved strain measurement of the net thrust for the cases of air only and hydrogen fuel injected into air at approximately 9 MJ/kg nozzle supply enthalpy. The gain in thrust due to combustion is visible in this result.

Tuttle, Sean↗
NASA NTRS1993

Axisymmetric computational fluid dynamics analysis of a film/dump-cooled rocket nozzle plume

Prediction of convective base heating rates for a new launch vehicle presents significant challenges to analysts concerned with base environments. The present effort seeks to augment classical base heating scaling techniques via a detailed investigation of the exhaust plume shear layer of a single H2/O2 Space Transportation Main Engine (STME). Use of fuel-rich turbine exhaust to cool the STME nozzle presented concerns regarding potential recirculation of these gases to the base region with attendant increase in the base heating rate. A pressure-based full Navier-Stokes computational fluid dynamics (CFD) code with finite rate chemistry is used to predict plumes for vehicle altitudes of 10 kft and 50 kft. Levels of combustible species within the plume shear layers are calculated in order to assess assumptions made in the base heating analysis.

Tucker, P. K.↗
NASA NTRS2012

Dual-Mode Scramjet Flameholding Operability Measurements

Flameholding measurements were made in two different direct connect combustor facilities that were designed to simulate a cavity flameholder in the flowfield of a hydrocarbon fueled dual-mode scramjet combustor. The presence of a shocktrain upstream of the flameholder has a significant impact on the inlet flow to the combustor and on the flameholding limits. A throttle was installed in the downstream end of the test rigs to provide the needed back-pressurization and decouple the operation of the flameholder from the backpressure formed by heat release and thermal choking, as in a flight engine. Measurements were made primarily with ethylene fuel but a limited number of tests were also performed with heated gaseous JP-7 fuel injection. The flameholding limits were measured by ramping inlet air temperature down until blowout was observed. The tests performed in the United Technologies Research Center (UTRC) facility used a hydrogen fueled vitiated air heater, Mach 2.2 and 3.3 inlet nozzles, a scramjet combustor rig with a 1.666 by 6 inch inlet and a 0.65 inch deep cavity. Mean blowout temperature measured at the baseline condition with ethylene fuel, the Mach 2.2 inlet and a cavity pressure of 21 psia was 1502 oR. Flameholding sensitivity to a variety of parameters was assessed. Blowout temperature was found to be most sensitive to fuel injection location and fuel flowrates and surprisingly insensitive to operating pressure (by varying both back-pressurization and inlet flowrate) and inlet Mach number. Video imaging through both the bottom and side wall windows was collected simultaneously and showed that the flame structure was quite unsteady with significant lateral movements as well as movement upstream of the flameholder. Experiments in the University of Virginia (UVa) test facility used a Mach 2 inlet nozzle with a 1 inch by 1.5 inch exit cross section, an aspect ratio of 1.5 versus 3.6 in the UTRC facility. The UVa facility tests were designed to measure the sensitivity of flameholding limits to inlet air vitiation by using electrically heated air and adding steam at levels to simulate vitiated air heaters. The measurements showed no significant difference in blowout temperature with inlet air mole fractions of steam from 0 to 6.7%.

airbreathing engines aircraft↗
NASA NTRS1959

Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

Messing, Wesley E.↗
NASA NTRS1988

Fluid spray simulation with two-fluid nozzles

Two-phase interacting flow inside a two-fluid fuel atomizer was investigated and a correction of aerodynamic and liquid-surface forces with characteristic drop diameter was obtained for liquid-jet breakup in Mach 1 gas flow. Nitrogen gas mass-flux was varied from 6 to 50 g/sq cm sec by using four differently sized two-fluid atomizers with nozzle diameters varyig from 0.32 to 0.56 cm. The correlation was derived by using the acoustic gas velocity, V sub c, as a basic parameter in defining and evaluating the dimensionless product of the Weber (We) and Reynolds (Re) numbers. By using the definition of WeRe, it was found that the ratio of orifice diameter to Sauter mean drop diameter could be correlated with the dimensionless ratio WeRe and the gas to liquid density ratio.

Ingebo, Robert D.↗
NASA NTRSDate not supplied

TPSAS-NF1676L-32577-DND

The ground testing of hypersonic vehicles requires the replication of the flow conditions at high speed and altitude. To produce these flow conditions, stored compressed air is expanded through a convergent-divergent nozzle to produce the required pressures and speed. However, the expansion process also cools the air, and so to match the flight temperatures, some form of heating needs to be employed. Vitiated heating is a common method, where the air is heated by burning fuel in the low speed side of the nozzle, and then oxygen is added to bring the level of O2 up to match that of standard air. The resulting flow no longer has the same composition as standard air, (except for the oxygen level), but rather contains combustion products, often referred to as vitiates as part of the mixture. The presence of these vitiates in the flow effects the wind tunnel test conditions. For example, the specific heat and molecular density are changed, the kinetics of combustion are altered and the flow features in the test article are altered. In this talk, the effects of the vitiates on the testing conditions and on the performance of test articles is described. Specific examples include the effects of molecular density on the flow, how combustion processes are altered by the presence of vitiates and how condensation of vitiates can alter the test conditions. In addition, the effect of the less common vitiates produced by arc-heated tunnels are mentioned, such as NO, non-thermal equilibrium and the presence of copper solids in the flow. Finally some thoughts on how large an effect these vitiates have on the test results are presented.

Andrew T Norris↗
NASA NTRS1983

Effects of tripropellant engines on earth-to-orbit vehicles

The effects of tripropellant engines on earth-to-orbit vehicles is examined in terms of their impact on engine configurations and launch capabilities. Hydrocarbon fuel with some oxygen is used in tripropellant fuels, with hydrogen as a back-up fluid for coolant and driving the pumps. Engine concepts which implement tripropellant fuels include a hydrogen-gas generator engine with staged combustion, a two-mode engine burning hydrocarbon fuel in one chamber and hydrogen in another, a dual expander engine with a central hydrocarbon nozzle and an annular hydrogen nozzle, and a dual throat engine. The use of hydrocarbons reduces the fuel weight by providing a higher specific impulse than with LOX-LH2 systems alone. Studies have shown that single-stage-to-orbit vehicles capable of lifting 13.6 Mg are possible with tripropellant engines. A dual-expander engine, is identified as offering the most fuel dry mass reduction for a given payload if cooling requirements can be satisfied. Further development of the tripropellant engines is concluded to be beneficial.

Martin, J. A.↗
NASA NTRS2009

Design, Activation, and Operation of the J2-X Subscale Simulator (JSS)

The purpose of this paper is to give a detailed description of the design, activation, and operation of the J2-X Subscale Simulator (JSS) installed in Cell 1 of the E3 test facility at Stennis Space Center, MS (SSC). The primary purpose of the JSS is to simulate the installation of the J2-X engine in the A3 Subscale Rocket Altitude Test Facility at SSC. The JSS is designed to give aerodynamically and thermodynamically similar plume properties as the J2-X engine currently under development for use as the upper stage engine on the ARES I and ARES V spacecraft. The JSS is a scale pressure fed, LOX/GH fueled rocket that is geometrically similar to the J2-X from the throat to the nozzle exit plane (NEP) and is operated at the same oxidizer to fuel ratios and chamber pressures. This paper describes the heritage hardware used as the basis of the JSS design, the newly designed rocket hardware, igniter systems used, and the activation and operation of the JSS.

Saunders, Grady P.↗