Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Cramer, J. M.; Greene, M. U.; Pal, S.; Santoro, R. J.; Turner, Jim (Technical Monitor)

2002-01-01

This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.

An Ejector Air Intake Design Method for a Novel Rocket-Based Combined-Cycle Rocket Nozzle

NASA Astrophysics Data System (ADS)

Waung, Timothy S.

Rocket-based combined-cycle (RBCC) vehicles have the potential to reduce launch costs through the use of several different air breathing engine cycles, which reduce fuel consumption. The rocket-ejector cycle, in which air is entrained into an ejector section by the rocket exhaust, is used at flight speeds below Mach 2. This thesis develops a design method for an air intake geometry around a novel RBCC rocket nozzle design for the rocket-ejector engine cycle. This design method consists of a geometry creation step in which a three-dimensional intake geometry is generated, and a simple flow analysis step which predicts the air intake mass flow rate. The air intake geometry is created using the rocket nozzle geometry and eight primary input parameters. The input parameters are selected to give the user significant control over the air intake shape. The flow analysis step uses an inviscid panel method and an integral boundary layer method to estimate the air mass flow rate through the intake geometry. Intake mass flow rate is used as a performance metric since it directly affects the amount of thrust a rocket-ejector can produce. The design method results for the air intake operating at several different points along the subsonic portion of the Ariane 4 flight profile are found to under predict mass flow rate by up to 8.6% when compared to three-dimensional computational fluid dynamics simulations for the same air intake.

RBCC Mixing Studies: Ejector Ramjet Design Optimization

NASA Technical Reports Server (NTRS)

1999-01-01

The research project reported herein extended over a period from October 1997 through August 1999. The research resulted in three technical papers presented at the AIAA/SAE/ASME/ASEE 35th Joint Propulsion Conference in Los Angeles in July 1999. These three papers are attached to this Executive Summary to constitute the final report. Objective: The objective of this research was to determine the mixing characteristics between the primary rocket jets and the turbine exhaust stream in a simulated Rocket Based Combined Cycle propulsion concept operating in the air augmented rocket mode.

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

Pumping Performance or RBCC Engine under Sea Level Static Condition

NASA Astrophysics Data System (ADS)

Kouchi, Toshinori; Tomioka, Sadatake; Kanda, Takeshi

Numerical simulations were conducted to predict the ejector pumping performance of a rocket-ramjet combined-cycle engine under a take-off condition. The numerical simulations revealed that the suction airflow was chocked at the exit of the engine throat when the ejector rocket was driven by cold N2 gas at the chamber pressure of 3MPa. When the ejector-driving gas was changed from cold N2 gas to hot combustion gas, the suction performance decreased remarkably. Mach contours in the engine revealed that the rocket plume constricted when the driving gas was the hot combustion gas. The change of the area of the stream tube area seemed to induce the pressure rise in the duct and decreasing in the pumping performance.

Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed

NASA Technical Reports Server (NTRS)

Yungster, S.; Trefny, C. J.

1999-01-01

An analysis of the Independent Ramjet Stream (IRS) cycle is presented. The IRS cycle is a variation of the conventional ejector-Ramjet, and is used at low speed in a rocket-based combined-cycle (RBCC) propulsion system. In this new cycle, complete mixing between the rocket and ramjet streams is not required, and a single rocket chamber can be used without a long mixing duct. Furthermore, this concept allows flexibility in controlling the thermal choke process. The resulting propulsion system is intended to be simpler, more robust, and lighter than an ejector-ramjet. The performance characteristics of the IRS cycle are analyzed for a new single-stage-to-orbit (SSTO) launch vehicle concept, known as "Trailblazer." The study is based on a quasi-one-dimensional model of the rocket and air streams at speeds ranging from lift-off to Mach 3. The numerical formulation is described in detail. A performance comparison between the IRS and ejector-ramjet cycles is also presented.

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

NASA Technical Reports Server (NTRS)

Landrum, D. Brian; Thames, Mignon; Parkinson, Doug; Gautney, Serena; Hawk, Clark

1999-01-01

Several tests were performed on a one-sixth scale Rocket Based Combined Cycle (RBCC) engine model at the University of Alabama in Huntsville. The UAH RBCC facility consists of a rectangular duct with a vertical strut mounted in the center. The scaled strut consists of two supersonic rocket nozzles with an embedded vertical turbine between the rocket nozzles. The tests included mass flow, flow visualization and horizontal pressure traverses. The mass flow test indicated a c:hoked condition when the rocket chamber pressure is between 200 psi and 300 psi. The flow visualization tests narrowed the rocket chamber pressure range from, 250 psi to 300 psi. Also, from this t.est, an assumption of a minimum

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions. Freejet tests of a candidate flowpath for this RBCC engine were conducted at the NASA Lewis Research Center's Hypersonic Tunnel Facility between July and September 1996. This paper describes the engine flowpath and installation, outlines the primary objectives of the program, and describes the overall results of this activity. Through this program 15 full duration tests, including 13 fueled tests were made. The first major achievement was the further demonstration of the HTF capability. The facility operated at conditions up to 1950 K and 7.34 MPa, simulating approximately Mach 6.6 flight. The initial tests were unfueled and focused on verifying both facility and engine starting. During these runs additional aerodynamic appliances were incorporated onto the facility diffuser to enhance starting. Both facility and engine starting were achieved. Further, the static pressure distributions compared well with the results previously obtained in a 40% subscale flowpath study conducted in the LERC 1X1 supersonic wind tunnel (SWT), as well as the results of CFD analysis. Fueled performance results were obtained for the engine at both simulated Mach 6 (1670 K) and Mach 6.6 (1950 K) conditions. For all these tests the primary fuel was liquid JP-10 with gaseous silane (a mixture of 20% SiH4 and 80% H2 by volume) as an ignitor/pilot. These tests verified performance of this engine flowpath in a freejet mode. High combustor pressures were reached and significant changes in axial force were achieved due to combustion. Future test plans include redistributing the fuel to improve mixing, and consequently performance, at higher equivalence ratios.

Performance of a RBCC Engine in Rocket-Operation

NASA Astrophysics Data System (ADS)

Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2003-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2002-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Rocket-based combined-cycle (RBCC) powered spaceliner class vehicle can advantageously employ vertical takeoff and landing (VTOL)

NASA Technical Reports Server (NTRS)

Escher, William J. D.

1995-01-01

The subject is next generation orbital space transporation, taken to be fully reusable non-staged 'aircraft like' systems targeted for routine, affordable access to space. Specifically, the takeoff and landing approach to be selected for such systems is considered, mainly from a propulsion viewpoint. Conventional wisdom has it that any transatmospheric-class vehicle which uses high-speed airbreathing propulsion modes (e.g., scramjet) intrinsically must utilize horizontal takeoff and landing, HTOHL. Although this may be true for all-airbreathing propulsion (i.e., no rocket content as in turboramjet propulsion), that emerging class of powerplant which integrally combines airbreathing and rocket propulsion, referred to as rocket-based combined-cycle (RBCC) propulsion, is considerably more flexible with respect to selecting takeoff/landing modes. In fact, it is proposed that any of the modes of interest may potentially be selected: HTOHL, VTOHL, VTOVL. To illustrate this surmise, the case of a previously documented RBCC-powered 'Spaceliner' class space transport concept, which is designed for vertical takeoff and landing, is examined. The 'RBCC' and 'Spaceliner' categories are first described for background. Departing form an often presumed HTOHL baseline, the leading design and operational advantages of moving to VTOVL are then elucidated. Technical substantiation that the RBCC approach, in fact, enables this capability (but also that of HTOHL and VTOVL) is provided, with extensive reference to case-in-point supporting studies. The paper closes with a set of conditional surmises bearing on its set of conclusions, which point up the operational cost advantages associated with selecting the vertical takeoff and landing mode combination (VTOL), uniquely offered by RBCC propulsion.

Rocket-based combined-cycle (RBCC) powered spaceliner class vehicle can advantageously employ vertical takeoff and landing (VTOL)

NASA Astrophysics Data System (ADS)

Escher, William J. D.

The subject is next generation orbital space transporation, taken to be fully reusable non-staged 'aircraft like' systems targeted for routine, affordable access to space. Specifically, the takeoff and landing approach to be selected for such systems is considered, mainly from a propulsion viewpoint. Conventional wisdom has it that any transatmospheric-class vehicle which uses high-speed airbreathing propulsion modes (e.g., scramjet) intrinsically must utilize horizontal takeoff and landing, HTOHL. Although this may be true for all-airbreathing propulsion (i.e., no rocket content as in turboramjet propulsion), that emerging class of powerplant which integrally combines airbreathing and rocket propulsion, referred to as rocket-based combined-cycle (RBCC) propulsion, is considerably more flexible with respect to selecting takeoff/landing modes. In fact, it is proposed that any of the modes of interest may potentially be selected: HTOHL, VTOHL, VTOVL. To illustrate this surmise, the case of a previously documented RBCC-powered 'Spaceliner' class space transport concept, which is designed for vertical takeoff and landing, is examined. The 'RBCC' and 'Spaceliner' categories are first described for background. Departing form an often presumed HTOHL baseline, the leading design and operational advantages of moving to VTOVL are then elucidated. Technical substantiation that the RBCC approach, in fact, enables this capability (but also that of HTOHL and VTOVL) is provided, with extensive reference to case-in-point supporting studies. The paper closes with a set of conditional surmises bearing on its set of conclusions, which point up the operational cost advantages associated with selecting the vertical takeoff and landing mode combination (VTOL), uniquely offered by RBCC propulsion.

The History and Promise of Combined Cycle Engines for Access to Space Applications

NASA Technical Reports Server (NTRS)

Clark, Casie

2010-01-01

For the summer of 2010, I have been working in the Aerodynamics and Propulsion Branch at NASA Dryden Flight Research Center studying combined-cycle engines, a high speed propulsion concept. Combined cycle engines integrate multiple propulsion systems into a single engine capable of running in multiple modes. These different modes allow the engine to be extremely versatile and efficient in varied flight conditions. The two most common types of combined cycle engines are Rocket-Based Combined Cycle (RBCC) and Turbine Based Combined Cycle (TBCC). The RBCC essentially combines a rocket and ramjet engine, while the TBCC integrates a turbojet and ramjet1. These two engines are able to switch between different propulsion modes to achieve maximum performance. Extensive conceptual and ground test studies of RBCC engines have been undertaken; however, an RBCC engine has never, to my knowledge, been demonstrated in flight. RBCC engines are of particular interest because they could potentially power a reusable launch vehicle (RLV) into space. The TBCC has been flight tested and shown to be effective at reaching supersonic speeds, most notably in the SR-71 Blackbird2.

A US History of Airbreathing/Rocket Combined-Cycle (RBCC) Propulsion for Powering Future Aerospace Transports, with a Look Ahead to the Year 2020

NASA Technical Reports Server (NTRS)

Escher, William J. D.

1999-01-01

A technohistorical and forward-planning overview of U.S. developments in combined airbreathing/rocket propulsion for advanced aerospace vehicle applications is presented. Such system approaches fall into one of two categories: (1) Combination propulsion systems (separate, non-interacting engines installed), and (2) Combined-Cycle systems. The latter, and main subject, comprises a large family of closely integrated engine types, made up of both airbreathing and rocket derived subsystem hardware. A single vehicle-integrated, multimode engine results, one capable of operating efficiently over a very wide speed and altitude range, atmospherically and in space. While numerous combination propulsion systems have reached operational flight service, combined-cycle propulsion development, initiated ca. 1960, remains at the subscale ground-test engine level of development. However, going beyond combination systems, combined-cycle propulsion potentially offers a compelling set of new and unique capabilities. These capabilities are seen as enabling ones for the evolution of Spaceliner class aerospace transportation systems. The following combined-cycle hypersonic engine developments are reviewed: (1) RENE (rocket engine nozzle ejector), (2) Cryojet and LACE, (3) Ejector Ramjet and its derivatives, (4) the seminal NASA NAS7-377 study, (5) Air Force/Marquardt Hypersonic Ramjet, (6) Air Force/Lockheed-Marquardt Incremental Scramjet flight-test project, (7) NASA/Garrett Hypersonic Research Engine (HRE), (8) National Aero-Space Plane (NASP), (9) all past projects; and such current and planned efforts as (10) the NASA ASTP-ART RBCC project, (11) joint CIAM/NASA DNSCRAM flight test,(12) Hyper-X, (13) Trailblazer,( 14) W-Vehicle and (15) Spaceliner 100. Forward planning programmatic incentives, and the estimated timing for an operational Spaceliner powered by combined-cycle engines are discussed.

A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

NASA Technical Reports Server (NTRS)

Stanley, Thomas Troy; Alexander, Reginald

1999-01-01

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

NASA Technical Reports Server (NTRS)

Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

2002-01-01

This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

Rocketdyne RBCC Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratckin, G.; Goldman, A.; Ortwerth, P.; Weisberg, S.

1999-01-01

Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

Rocketdyne Development of RBCC Engine for Low Cost Access to Space

NASA Technical Reports Server (NTRS)

Ortwerth, P.; Ratekin, G.; Goldman, A.; Emanuel, M.; Ketchum, A.; Horn, M.

1997-01-01

Rocketdyne is pursuing the conceptual design and development of a Rocket Based Combined Cycle (RBCC) engine for booster and SSTO, advanced reusable space transportation ARTT systems under contract with NASA Marshall Space Flight Center. The Rocketdyne concept is fixed geometry integrated Rocket, Ram Scramjet which is Hydrogen fueled and uses Hydrogen regenerative cooling. Vision vehicle integration studies have determined that scramjet operation to Mach 12 has high payoff for low cost reusable space transportation. Rocketdyne is internally developing versions of the concept for other applications in high speed aircraft and missiles with Hydrocarbon fuel systems. Subscale engine ground testing is underway for all modes of operation from takeoff to Mach 8. High altitude Rocket only mode tests will be completed as part of the ground test program to validate high expansion ratio performance. A unique feature of the ground test series is the inclusion of dynamic trajectory simulation with real time Mach number, altitude, engine throttling, and RBCC mode changes in a specially modified freejet test facility at GASL. Preliminary cold flow Air Augmented Rocket mode test results and Short Combustor tests have met program goals and have been used to integrate all modes of operation in a single combustor design with a fixed geometry inlet for design confirmation tests. A water cooled subscale engine is being fabricated and installed for test beginning the last quarter of 1997.

Investigation of Compressibility Effect for Aeropropulsive Shear Flows

NASA Technical Reports Server (NTRS)

Balasubramanyam, M. S.; Chen, C. P.

2005-01-01

Rocket Based Combined Cycle (RBCC) engines operate within a wide range of Mach numbers and altitudes. Fundamental fluid dynamic mechanisms involve complex choking, mass entrainment, stream mixing and wall interactions. The Propulsion Research Center at the University of Alabama in Huntsville is involved in an on- going experimental and numerical modeling study of non-axisymmetric ejector-based combined cycle propulsion systems. This paper attempts to address the modeling issues related to mixing, shear layer/wall interaction in a supersonic Strutjet/ejector flow field. Reynolds Averaged Navier-Stokes (RANS) solutions incorporating turbulence models are sought and compared to experimental measurements to characterize detailed flow dynamics. The effect of compressibility on fluids mixing and wall interactions were investigated using an existing CFD methodology. The compressibility correction to conventional incompressible two- equation models is found to be necessary for the supersonic mixing aspect of the ejector flows based on 2-D simulation results. 3-D strut-base flows involving flow separations were also investigated.

Performance Evaluation of the NASA GTX RBCC Flowpath

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Palac, Donald T.; Trefny, Charles J.; Roche, Joseph M.

2001-01-01

The NASA Glenn Research Center serves as NASAs lead center for aeropropulsion. Several programs are underway to explore revolutionary airbreathing propulsion systems in response to the challenge of reducing the cost of space transportation. Concepts being investigated include rocket-based combined cycle (RBCC), pulse detonation wave, and turbine-based combined cycle (TBCC) engines. The GTX concept is a vertical launched, horizontal landing, single stage to orbit (SSTO) vehicle utilizing RBCC engines. The propulsion pod has a nearly half-axisymmetric flowpath that incorporates a rocket and ram-scramjet. The engine system operates from lift-off up to above Mach 10, at which point the airbreathing engine flowpath is closed off, and the rocket alone powers the vehicle to orbit. The paper presents an overview of the research efforts supporting the development of this RBCC propulsion system. The experimental efforts of this program consist of a series of test rigs. Each rig is focused on development and optimization of the flowpath over a specific operating mode of the engine. These rigs collectively establish propulsion system performance over all modes of operation, therefore, covering the entire speed range. Computational Fluid Mechanics (CFD) analysis is an important element of the GTX propulsion system development and validation. These efforts guide experiments and flowpath design, provide insight into experimental data, and extend results to conditions and scales not achievable in ground test facilities. Some examples of important CFD results are presented.

Propulsion/ASME Rocket-Based Combined Cycle Activities in the Advanced Space Transportation Program Office

NASA Technical Reports Server (NTRS)

Hueter, Uwe; Turner, James

1998-01-01

NASA's Office Of Aeronautics and Space Transportation Technology (OASTT) has establish three major coals. "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville,Ala. focuses on future space transportation technologies under the "Access to Space" pillar. The Advanced Reusable Technologies (ART) Project, part of ASTP, focuses on the reusable technologies beyond those being pursued by X-33. The main activity over the past two and a half years has been on advancing the rocket-based combined cycle (RBCC) technologies. In June of last year, activities for reusable launch vehicle (RLV) airframe and propulsion technologies were initiated. These activities focus primarily on those technologies that support the year 2000 decision to determine the path this country will take for Space Shuttle and RLV. In February of this year, additional technology efforts in the reusable technologies were awarded. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion. Aerojet, Boeing-Rocketdyne and Pratt & Whitney were selected for a two-year period to design, build and ground test their RBCC engine concepts. In addition, ASTROX, Pennsylvania State University (PSU) and University of Alabama in Huntsville also conducted supporting activities. The activity included ground testing of components (e.g., injectors, thrusters, ejectors and inlets) and integrated flowpaths. An area that has caused a large amount of difficulty in the testing efforts is the means of initiating the rocket combustion process. All three of the prime contractors above were using silane (SiH4) for ignition of the thrusters. This follows from the successful use of silane in the NASP program for scramjet ignition. However, difficulties were immediately encountered when silane (an 80/20 mixture of hydrogen/silane) was used for rocket ignition.

Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Chojnacki, Kent T.

1992-01-01

The goal of the Rocket-Based Combined Cycle (RBCC) Propulsion Technology Workshop was to assess the RBCC propulsion system's viability for Earth-to-Orbit (ETO) transportation systems. This was accomplished by creating a forum (workshop) in which past work in the field of RBCC propulsion systems was reviewed, current technology status was evaluated, and future technology programs in the field of RBCC propulsion systems were postulated, discussed, and recommended.

Summary of Rocketdyne Engine A5 Rocket Based Combined Cycle Testing

NASA Technical Reports Server (NTRS)

Ketchum. A.; Emanuel, Mark; Cramer, John

1998-01-01

Rocketdyne Propulsion and Power (RPP) has completed a highly successful experimental test program of an advanced rocket based combined cycle (RBCC) propulsion system. The test program was conducted as part of the Advanced Reusable Technology program directed by NASA-MSFC to demonstrate technologies for low-cost access to space. Testing was conducted in the new GASL Flight Acceleration Simulation Test (FAST) facility at sea level (Mach 0), Mach 3.0 - 4.0, and vacuum flight conditions. Significant achievements obtained during the test program include 1) demonstration of engine operation in air-augmented rocket mode (AAR), ramjet mode and rocket mode and 2) smooth transition from AAR to ramjet mode operation. Testing in the fourth mode (scramjet) is scheduled for November 1998.

A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

NASA Technical Reports Server (NTRS)

Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

2000-01-01

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC engine model. HYFIM performs the aerodynamic analysis of forebodies and inlet characteristics of RBCC powered SSTO launch vehicles. HYFIM is applicable to the analysis of the ramjet/scramjet engine operations modes (Mach 3-12), and provides estimates of parameters such as air capture area, shock-on-lip Mach number, design Mach number, compression ratio, etc., based on a basic geometry routine for modeling axisymmetric cones, 2-D wedge geometries. HYFIM also estimates the variation of shock layer properties normal to the forebody surface. The thermal protection system (TPS) is directly linked to determination of the vehicle moldline and the shaping of the trajectory. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. The need to analyze vehicle forebody and engine inlet is critical to be able to design the RBCC vehicle. To adequately determine insulation masses for an RBCC vehicle, the hypersonic aerodynamic environment and aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass. E-PSURBCC is an "engine performance" model and requires the specification of inlet air static temperature and pressure as well as Mach number (which it pulls from the HYFIM and POST trajectory files), and calculates the corresponding stagnation properties. The engine air flow path geometry includes inlet, a constant area section where the rocket is positioned, a subsonic diffuser, a constant area afterburner, and either a converging nozzle or a converging-diverging nozzle. The current capabilities of E-PSURBCC ejector and ramjet mode treatment indicated that various complex flow phenomena including multiple choking and internal shocks can occur for combinations of geometry/flow conditions. For a given input deck defining geometry/flow conditions, the program first goes through a series of checks to establish whether the input parameters are sound in terms of a solution path. If the vehicle/engine performance fails mission goals, the engineer is able to collaboratively alter the vehicle moldline to change aerodynamics, or trajectory, or some other input to achieve orbit. The problem described is an example of the need for collaborative design and analysis. RECIPE is a cross-platform application capable of hosting a number of engineers and designers across the Internet for distributed and collaborative engineering environments. Such integrated system design environments allow for collaborative team design analysis for performing individual or reduced team studies. To facilitate the larger number of potential runs that may need to be made, RECIPE connects the computer codes that calculate the trajectory data, aerodynamic data based on vehicle geometry, heat rate data, TPS masses, and vehicle and engine performance, so that the output from each tool is easily transferred to the model input files that need it.

Accomplishments of the Advanced Reusable Technologies (ART) RBCC Project at NASA/Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Nelson, Karl W.; McArthur, J. Craig (Technical Monitor)

2001-01-01

The focus of the NASA / Marshall Space Flight Center (MSFC) Advanced Reusable Technologies (ART) project is to advance and develop Rocket-Based Combined-Cycle (RBCC) technologies. The ART project began in 1996 as part of the Advanced Space Transportation Program (ASTP). The project is composed of several activities including RBCC engine ground testing, tool development, vehicle / mission studies, and component testing / development. The major contractors involved in the ART project are Aerojet and Rocketdyne. A large database of RBCC ground test data was generated for the air-augmented rocket (AAR), ramjet, scramjet, and ascent rocket modes of operation for both the Aerojet and Rocketdyne concepts. Transition between consecutive modes was also demonstrated as well as trajectory simulation. The Rocketdyne freejet tests were conducted at GASL in the Flight Acceleration Simulation Test (FAST) facility. During a single test, the FAST facility is capable of simulating both the enthalpy and aerodynamic conditions over a range of Mach numbers in a flight trajectory. Aerojet performed freejet testing in the Pebble Bed facility at GASL as well as direct-connect testing at GASL. Aerojet also performed sea-level static (SLS) testing at the Aerojet A-Zone facility in Sacramento, CA. Several flight-type flowpath components were developed under the ART project. Aerojet designed and fabricated ceramic scramjet injectors. The structural design of the injectors will be tested in a simulated scramjet environment where thermal effects and performance will be assessed. Rocketdyne will be replacing the cooled combustor in the A5 rig with a flight-weight combustor that is near completion. Aerojet's formed duct panel is currently being fabricated and will be tested in the SLS rig in Aerojet's A-Zone facility. Aerojet has already successfully tested a cooled cowl panel in the same facility. In addition to MSFC, other NASA centers have contributed to the ART project as well. Inlet testing and parametrics were performed at NASA / Glenn Research Center (GRC) and NASA / Langley Research Center (LaRC) for both the Aerojet and Rocketdyne concepts. LaRC conducted an Air-Breathing Launch Vehicle (ABLV) study for several vehicle concepts with RBCC propulsion systems. LaRC is also performing a CFD analysis of the ramjet mode for both flowpaths based on GASL test conditions. A study was performed in 1999 to investigate the feasibility of performing an RBCC flight test on the NASA / Dryden Flight Research Center (DFRC) SR-71 aircraft. Academia involvement in the ART project includes parametric RBCC flowpath testing by Pennsylvania State University (PSU). In addition to thrust and wall static pressure measurements, PSU is also using laser diagnostics to analyze the flowfield in the test rig. MSFC is performing CFD analysis of the PSU rig at select test conditions for model baseline and validation. Also, Georgia Institute of Technology (GT) conducted a vision vehicle study using the Aerojet RBCC concept. Overall, the ART project has been very successful in advancing RBCC technology. Along the way, several major milestones were achieved and "firsts" accomplished. For example, under the ART project, the first dynamic trajectory simulation testing was performed and the Rocketdyne engine A5 logged over one hour of accumulated test time. The next logical step is to develop and demonstrate a flight-weight RBCC engine system.

Development of a 12-Thrust Chamber Kerosene /Oxygen Primary Rocket Sub-System for an Early (1964) Air-Augmented Rocket Ground-Test System

NASA Technical Reports Server (NTRS)

Pryor, D.; Hyde, E. H.; Escher, W. J. D.

1999-01-01

Airbreathing/Rocket combined-cycle, and specifically rocket-based combined- cycle (RBCC), propulsion systems, typically employ an internal engine flow-path installed primary rocket subsystem. To achieve acceptably short mixing lengths in effecting the "air augmentation" process, a large rocket-exhaust/air interfacial mixing surface is needed. This leads, in some engine design concepts, to a "cluster" of small rocket units, suitably arrayed in the flowpath. To support an early (1964) subscale ground-test of a specific RBCC concept, such a 12-rocket cluster was developed by NASA's Marshall Space Flight Center (MSFC). The small primary rockets used in the cluster assembly were modified versions of an existing small kerosene/oxygen water-cooled rocket engine unit routinely tested at MSFC. Following individual thrust-chamber tests and overall subsystem qualification testing, the cluster assembly was installed at the U. S. Air Force's Arnold Engineering Development Center (AEDC) for RBCC systems testing. (The results of the special air-augmented rocket testing are not covered here.) While this project was eventually successfully completed, a number of hardware integration problems were met, leading to catastrophic thrust chamber failures. The principal "lessons learned" in conducting this early primary rocket subsystem experimental effort are documented here as a basic knowledge-base contribution for the benefit of today's RBCC research and development community.

Rocket Based Combined Cycle (RBCC) Engine

NASA Technical Reports Server (NTRS)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

The Strutjet Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

1998-01-01

The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine. RBCC engines exhibit a high potential for lowering the operating cost of launching payloads into orbit. Two sources of cost reductions can be identified. First, RBCC powered vehicles require only 20% takeoff thrust compared to conventional rockets, thereby lowering the thrust requirements and the replacement cost of the engines. Second, due to the higher structural and thermal margins achievable with RBCC engines coupled with a higher degree of subsystem redundance lower maintenance and operating cost are obtainable.

Research Technology (ASTP) Rocket Based Combined Cycle (RBCC) Engine

NASA Technical Reports Server (NTRS)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

Rocket Based Combined Cycle (RBCC) engine inlet

NASA Technical Reports Server (NTRS)

2004-01-01

Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

Rocket Based Combined Cycle (RBCC) Propulsion Workshop, volume 2

NASA Technical Reports Server (NTRS)

Chojnacki, Kent T.

1992-01-01

The goal of the Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop, was to impart technology information to the propulsion community with respect to hypersonic combined cycle propulsion capabilities. The major recommendation resulting from this technology workshop was as follows: conduct a systems-level applications study to define the desired propulsion system and vehicle technology requirements for LEO launch vehicles. All SSTO and TSTO options using the various propulsion systems (airbreathing combined cycle, rocket-based combined cycle, and all rocket) must be considered. Such a study should be accomplished as soon as possible. It must be conducted with a consistent set of ground rules and assumptions. Additionally, the study should be conducted before any major expenditures on a RBCC technology development program occur.

Numerical simulation of divergent rocket-based-combined-cycle performances under the flight condition of Mach 3

NASA Astrophysics Data System (ADS)

Cui, Peng; Xu, WanWu; Li, Qinglian

2018-01-01

Currently, the upper operating limit of the turbine engine is Mach 2+, and the lower limit of the dual-mode scramjet is Mach 4. Therefore no single power systems can operate within the range between Mach 2 + and Mach 4. By using ejector rockets, Rocket-based-combined-cycle can work well in the above scope. As the key component of Rocket-based-combined-cycle, the ejector rocket has significant influence on Rocket-based-combined-cycle performance. Research on the influence of rocket parameters on Rocket-based-combined-cycle in the speed range of Mach 2 + to Mach 4 is scarce. In the present study, influences of Mach number and total pressure of the ejector rocket on Rocket-based-combined-cycle were analyzed numerically. Due to the significant effects of the flight conditions and the Rocket-based-combined-cycle configuration on Rocket-based-combined-cycle performances, flight altitude, flight Mach number, and divergence ratio were also considered. The simulation results indicate that matching lower altitude with higher flight Mach numbers can increase Rocket-based-combined-cycle thrust. For another thing, with an increase of the divergent ratio, the effect of the divergent configuration will strengthen and there is a limit on the divergent ratio. When the divergent ratio is greater than the limit, the effect of divergent configuration will gradually exceed that of combustion on supersonic flows. Further increases in the divergent ratio will decrease Rocket-based-combined-cycle thrust.

XCALIBUR: a Vertical Takeoff TSTO RLV Concept with a HEDM Upperstage and a Scram-Rocket Booster

NASA Astrophysics Data System (ADS)

Bradford, J.

2002-01-01

A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15 staging condition. The upperstage is released from the booster and carries the 6,820 kg of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro- mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). The booster system is enabled main propulsion system which utilizes four RBCC engines. These engines operate in four distinct modes: air- augmented rocket (AAR), ramjet, scram-rocket, and all-rocket. The booster operates in AAR mode from takeoff to Mach 3, with ramjet mode operation from Mach 3 to Mach 6. The rocket re-ignition for scram-rocket mode occurs at Mach 6, with all-rocket mode from Mach 14 to the staging condition. The extended utilization of the scram-rocket mode greatly improves vehicle performance by providing superior vehicle acceleration when compared to the scramjet mode performance over the same flight region. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable maximum dynamic pressure. A complete vehicle system life-cycle analysis was performed in an automated, multi-disciplinary design environment. Automated disciplinary performance analysis tools include: trajectory (POST), propulsion (SCCREAM), aeroheating (TCAT II), and an Excel spreadsheet for component weight estimation. These tools were automated using `file wrappers' in Phoenix Integration's ModelCenter collaborative design environment. Performance tools utilized for the analysis, but not requiring automation included IDEAS for solid modeling and APAS for the aerodynamic analysis. The paper describes the vehicle concept and operation, discussing the types of technologies used and the nominal flight scenario. A brief discussion explaining the decision-making process for the vehicle configuration is included. For cost predictions, NAFCOM-derived cost estimating relationships were used. Economic predictions were developed using a number of codes, including CABAM (financials), AATe (operations), and GTSafetyII (safety and reliability).

Highly Reusable Space Transportation (HRST) Baseline Concepts and Analysis: Rocket/RBCC Options. Part 2; A Comparative Study

NASA Technical Reports Server (NTRS)

Woodcock, Gordon

1997-01-01

This study is an extension of a previous effort by the Principal Investigator to develop baseline data to support comparative analysis of Highly Reusable Space Transportation (HRST) concepts. The analyses presented herin develop baseline data bases for two two-stage-to-orbit (TSTO) concepts: (1) Assisted horizontal take-off all rocket (assisted HTOHL); and (2) Assisted vertical take-off rocket based combined cycle (RBCC). The study objectives were to: (1) Provide configuration definitions and illustrations for assisted HTOHL and assisted RBCC; (2) Develop a rationalization approach and compare these concepts with the HRST reference; and (3) Analyze TSTO configurations which try to maintain SSTO benefits while reducing inert weight sensitivity.

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 10 rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Mixing of Supersonic Jets in a RBCC Strutjet Propulsion System

NASA Technical Reports Server (NTRS)

Muller, S.; Hawk, Clark W.; Bakker, P. G.; Parkinson, D.; Turner, M.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during take-off and low speed flight. A scale model of the Strutjet device was built and tested to investigate the mixing of the streams as a function of distance from the Strut exit plane in simulated sea level take-off conditions. The Planar Laser Induced Fluorescence (PLIF) diagnostic method has been employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air. The ratio of the pressure in the turbine exhaust to that in the rocket nozzle wall at the point where the two jets meet, is the independent variable in these experiments. Tests were accomplished at values of 1.0 (the original design point), 1.5 and 2.0 for this parameter at 8 locations downstream of the rocket nozzle exit. The results illustrate the development of the mixing zone from the exit plane of the strut to a distance of about 18 equivalent rocket nozzle exit diameters downstream (18"). These images show the turbine exhaust to be confined until a short distance downstream. The expansion into the ingested air is more pronounced at a pressure ratio of 1.0 and 1.5 and shows that mixing with this air would likely begin at a distance of 2" downstream of the nozzle exit plane. Of the pressure ratios tested in this research, 2.0 is the best value for delaying the mixing at the operating conditions considered.

Computational Analysis of the Combustion Processes in an Axisymmetric, RBCC Flowpath

NASA Technical Reports Server (NTRS)

Steffen, Christopher J., Jr.; Yungster, Shaye

2001-01-01

Computational fluid dynamic simulations have been used to study the combustion processes within an axisymmetric, RBCC flowpath. Two distinct operating modes have been analyzed to date, including the independent ramjet stream (IRS) cycle and the supersonic combustion ramjet (scramJet) cycle. The IRS cycle investigation examined the influence of fuel-air ratio, fuel distribution, and rocket chamber pressure upon the combustion physics and thermal choke characteristics. Results indicate that adjustment of the amount and radial distribution of fuel can control the thermal choke point. The secondary massflow rate was very sensitive to the fuel-air ratio and the rocket chamber pressure. The scramjet investigation examined the influence of fuel-air ratio and fuel injection schedule upon combustion performance estimates. An analysis of the mesh-dependence of these calculations was presented. Jet penetration data was extracted from the three-dimensional simulations and compared favorably with experimental correlations of similar flows. Results indicate that combustion efficiency was very sensitive to the fuel schedule.

Design and Fabrication of the ISTAR Direct-Connect Combustor Experiment at the NASA Hypersonic Tunnel Facility

NASA Technical Reports Server (NTRS)

Lee, Jin-Ho; Krivanek, Thomas M.

2005-01-01

The Integrated Systems Test of an Airbreathing Rocket (ISTAR) project was a flight demonstration project initiated to advance the state of the art in Rocket Based Combined Cycle (RBCC) propulsion development. The primary objective of the ISTAR project was to develop a reusable air breathing vehicle and enabling technologies. This concept incorporated a RBCC propulsion system to enable the vehicle to be air dropped at Mach 0.7 and accelerated up to Mach 7 flight culminating in a demonstration of hydrocarbon scramjet operation. A series of component experiments was planned to reduce the level of risk and to advance the technology base. This paper summarizes the status of a full scale direct connect combustor experiment with heated endothermic hydrocarbon fuels. This is the first use of the NASA GRC Hypersonic Tunnel facility to support a direct-connect test. The technical and mechanical challenges involved with adapting this facility, previously used only in the free-jet configuration, for use in direct connect mode will be also described.

Rocket-Based Combined Cycle Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

2001-01-01

The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

Structural Design of Dual-Mode Ramjets and Associated System Issues (Conception structurale des statoreacteurs mixtes et defis systeme associes)

DTIC Science & Technology

2010-09-01

Institute PAO Protection Against Oxidation RBCC Rocket Based Combined Cycle RSL Reusable Space Launchers SSTO Single Stage to Orbit TOW Take-Off Weight...Orbit) or Single Stage To Orbit ( SSTO ) vehicles, or other kind of hypersonic vehicles. For example, in the scope of the French PREPHA program, the...study of a generic SSTO vehicle led to conclusion that the best type of airbreathing engine could be the dual-mode ramjet (subsonic then supersonic

Oxidizer Selection for the ISTAR Program (Liquid Oxygen versus Hydrogen Peroxide)

NASA Technical Reports Server (NTRS)

Quinn, Jason Eugene; Koelbl, Mary E. (Technical Monitor)

2002-01-01

This paper discusses a study of two alternate oxidizers, liquid oxygen and hydrogen peroxide, for use in a rocket based combined cycle (RBCC) demonstrator vehicle. The flight vehicle is baselined as an airlaunched self-powered Mach 0.7 to 7 demonstration of an RBCC engine through all or its air breathing propulsion modes. Selection of an alternate oxidizer has the potential to lower overall vehicle size, system complexity/ cost and ultimately the total program risk. This trade study examined the oxidizer selection effects upon the overall vehicle performance, safety and operations. After consideration of all the technical and programmatic details available at this time, 90% hydrogen peroxide was selected over liquid oxygen for use in this program.

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 19 equivalent rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Rocket-Based Combined-Cycle (RBCC) Propulsion Technology Workshop. Tutorial session

NASA Technical Reports Server (NTRS)

1992-01-01

The goal of this workshop was to illuminate the nation's space transportation and propulsion engineering community on the potential of hypersonic combined cycle (airbreathing/rocket) propulsion systems for future space transportation applications. Four general topics were examined: (1) selections from the expansive advanced propulsion archival resource; (2) related propulsion systems technical backgrounds; (3) RBCC engine multimode operations related subsystem background; and (4) focused review of propulsion aspects of current related programs.

ASTP RBCC Activities

NASA Technical Reports Server (NTRS)

Nelson, Karl W.; McArthur, Craig; Leopard, Larry (Technical Monitor)

2000-01-01

This presentation reviews the activities of the Advanced Space Transportation Program (ASTP) in the development of Rocket-Based Combined Cycle (RBCC)technology. The document consist of the presentation slides for a talk scheduled to be given to the World Aviation Congress and Exhibit of SAE. Included in the review is discussion of recent accomplishments in the area of Advanced Reusable technologies (ART), which includes work in flowpath testing, and system studies of the various vehicle/engine combinations including RBCC, Turbine Based Combined Cycle (TBCC) and Pulsed Detonation Engine (PDE). Pictures of the proposed RBCC Flowpaths are included. The next steps in the development process are reviewed.

JANNAF 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee Joint Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

1999-01-01

Volume 1, the first of three volumes is a compilation of 16 unclassified/unlimited-technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee held jointly with the 181 Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered include overviews of RBCC and PDE hypersonic technology, Hyper-X propulsion ground testing, development of JP-8 for hypersonic vehicle applications, numerical simulation of dual-mode SJ combustion, V&V of M&S computer codes, MHD SJ and Rocket Based Combined Cycle (RBCC) launch vehicle concepts, and Pulse Detonation Engine (PDE) propulsion technology development including fundamental investigations, modeling, aerodynamics, operation and performance.

Reuse fo a Cold War Surveillance Drone to Flight Test a NASA Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Brown, T. M.; Smith, Norm

1999-01-01

Plans for and early feasibility investigations into the modification of a Lockheed D21B drone to flight test the DRACO Rocket Based Combined Cycle (RBCC) engine are discussed. Modifications include the addition of oxidizer tanks, modern avionics systems, actuators, and a vehicle recovery system. Current study results indicate that the D21B is a suitable candidate for this application and will allow demonstrations of all DRACO engine operating modes at Mach numbers between 0.8 and 4.0. Higher Mach numbers may be achieved with more extensive modification. Possible project risks include low speed stability and control, and recovery techniques.

Multidisciplinary design of a rocket-based combined cycle SSTO launch vehicle using Taguchi methods

NASA Technical Reports Server (NTRS)

Olds, John R.; Walberg, Gerald D.

1993-01-01

Results are presented from the optimization process of a winged-cone configuration SSTO launch vehicle that employs a rocket-based ejector/ramjet/scramjet/rocket operational mode variable-cycle engine. The Taguchi multidisciplinary parametric-design method was used to evaluate the effects of simultaneously changing a total of eight design variables, rather than changing them one at a time as in conventional tradeoff studies. A combination of design variables was in this way identified which yields very attractive vehicle dry and gross weights.

A Thermal Management Systems Model for the NASA GTX RBCC Concept

NASA Technical Reports Server (NTRS)

Traci, Richard M.; Farr, John L., Jr.; Laganelli, Tony; Walker, James (Technical Monitor)

2002-01-01

The Vehicle Integrated Thermal Management Analysis Code (VITMAC) was further developed to aid the analysis, design, and optimization of propellant and thermal management concepts for advanced propulsion systems. The computational tool is based on engineering level principles and models. A graphical user interface (GUI) provides a simple and straightforward method to assess and evaluate multiple concepts before undertaking more rigorous analysis of candidate systems. The tool incorporates the Chemical Equilibrium and Applications (CEA) program and the RJPA code to permit heat transfer analysis of both rocket and air breathing propulsion systems. Key parts of the code have been validated with experimental data. The tool was specifically tailored to analyze rocket-based combined-cycle (RBCC) propulsion systems being considered for space transportation applications. This report describes the computational tool and its development and verification for NASA GTX RBCC propulsion system applications.

Development of an Integrated Nozzle for a Symmetric, RBCC Launch Vehicle Configuration

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Canabal, Francisco, III; Rice, Tharen; Blaha, Bernard

2000-01-01

The development of rocket based combined cycle (RBCC) engines is highly dependent upon integrating several different modes of operation into a single system. One of the key components to develop acceptable performance levels through each mode of operation is the nozzle. It must be highly integrated to serve the expansion processes of both rocket and air-breathing modes without undue weight, drag, or complexity. The NASA GTX configuration requires a fixed geometry, altitude-compensating nozzle configuration. The initial configuration, used mainly to estimate weight and cooling requirements was a 1 So half-angle cone, which cuts a concave surface from a point within the flowpath to the vehicle trailing edge. Results of 3-D CFD calculations on this geometry are presented. To address the critical issues associated with integrated, fixed geometry, multimode nozzle development, the GTX team has initiated a series of tasks to evolve the nozzle design, and validate performance levels. An overview of these tasks is given. The first element is a design activity to develop tools for integration of efficient expansion surfaces With the existing flowpath and vehicle aft-body, and to develop a second-generation nozzle design. A preliminary result using a "streamline-tracing" technique is presented. As the nozzle design evolves, a combination of 3-D CFD analysis and experimental evaluation will be used to validate the design procedure and determine the installed performance for propulsion cycle modeling. The initial experimental effort will consist of cold-flow experiments designed to validate the general trends of the streamline-tracing methodology and anchor the CFD analysis. Experiments will also be conducted to simulate nozzle performance during each mode of operation. As the design matures, hot-fire tests will be conducted to refine performance estimates and anchor more sophisticated reacting-flow analysis.

Rocket-Induced Magnetohydrodynamic Ejector: A Single-Stage-to-Orbit Advanced Propulsion Concept

NASA Technical Reports Server (NTRS)

Cole, John; Campbell, Jonathan; Robertson, Anthony

1995-01-01

During the atmospheric boost phase of a rocket trajectory, magnetohydrodynamic (MHD) principles can be utilized to augment the thrust by several hundred percent without the input of additional energy. The concept is an MHD implementation of a thermodynamic ejector. Some ejector history is described and some test data showing the impressive thrust augmentation capabilities of thermodynamic ejectors are provided. A momentum and energy balance is used to derive the equations to predict the MHD ejector performance. Results of these equations are compared with the test data and then applied to a specific performance example. The rocket-induced MHD ejector (RIME) engine is described and a status of the technology and availability of the engine components is provided. A top level vehicle sizing analysis is performed by scaling existing MHD designs to the required flight vehicle levels. The vehicle can achieve orbit using conservative technology. Modest improvements are suggested using recently developed technologies, such as superconducting magnets, which can improve predicted performance well beyond those expected for current single-stage-to-orbit (SSTO) designs.

Space Transportation Propulsion Systems

NASA Technical Reports Server (NTRS)

Liou, Meng-Sing; Stewart, Mark E.; Suresh, Ambady; Owen, A. Karl

2001-01-01

This report outlines the Space Transportation Propulsion Systems for the NPSS (Numerical Propulsion System Simulation) program. Topics include: 1) a review of Engine/Inlet Coupling Work; 2) Background/Organization of Space Transportation Initiative; 3) Synergy between High Performance Computing and Communications Program (HPCCP) and Advanced Space Transportation Program (ASTP); 4) Status of Space Transportation Effort, including planned deliverables for FY01-FY06, FY00 accomplishments (HPCCP Funded) and FY01 Major Milestones (HPCCP and ASTP); and 5) a review current technical efforts, including a review of the Rocket-Based Combined-Cycle (RBCC), Scope of Work, RBCC Concept Aerodynamic Analysis and RBCC Concept Multidisciplinary Analysis.

A Technology Pathway for Airbreathing, Combined-Cycle, Horizontal Space Launch Through SR-71 Based Trajectory Modeling

NASA Technical Reports Server (NTRS)

Kloesel, Kurt J.; Ratnayake, Nalin A.; Clark, Casie M.

2011-01-01

Access to space is in the early stages of commercialization. Private enterprises, mainly under direct or indirect subsidy by the government, have been making headway into the LEO launch systems infrastructure, of small-weight-class payloads of approximately 1000 lbs. These moderate gains have emboldened the launch industry and they are poised to move into the middle-weight class (roughly 5000 lbs). These commercially successful systems are based on relatively straightforward LOX-RP, two-stage, bi-propellant rocket technology developed by the government 40 years ago, accompanied by many technology improvements. In this paper we examine a known generic LOX-RP system with the focus on the booster stage (1st stage). The booster stage is then compared to modeled Rocket-Based and Turbine-Based Combined Cycle booster stages. The air-breathing propulsion stages are based on/or extrapolated from known performance parameters of ground tested RBCC (the Marquardt Ejector Ramjet) and TBCC (the SR-71/J-58 engine) data. Validated engine models using GECAT and SCCREAM are coupled with trajectory optimization and analysis in POST-II to explore viable launch scenarios using hypothetical aerospaceplane platform obeying the aerodynamic model of the SR-71. Finally, and assessment is made of the requisite research technology advances necessary for successful commercial and government adoption of combined-cycle engine systems for space access.

Preliminary Sizing of Vertical Take-off Rocket-based Combined-cycle Powered Launch Vehicles

NASA Technical Reports Server (NTRS)

Roche, Joseph M.; McCurdy, David R.

2001-01-01

The task of single-stage-to-orbit has been an elusive goal due to propulsion performance, materials limitations, and complex system integration. Glenn Research Center has begun to assemble a suite of relationships that tie Rocket-Based Combined-Cycle (RBCC) performance and advanced material data into a database for the purpose of preliminary sizing of RBCC-powered launch vehicles. To accomplish this, a near optimum aerodynamic and structural shape was established as a baseline. The program synthesizes a vehicle to meet the mission requirements, tabulates the results, and plots the derived shape. A discussion of the program architecture and an example application is discussed herein.

Three Dimensional Numerical Simulation of Rocket-based Combined-cycle Engine Response During Mode Transition Events

NASA Technical Reports Server (NTRS)

Edwards, Jack R.; McRae, D. Scott; Bond, Ryan B.; Steffan, Christopher (Technical Monitor)

2003-01-01

The GTX program at NASA Glenn Research Center is designed to develop a launch vehicle concept based on rocket-based combined-cycle (RBCC) propulsion. Experimental testing, cycle analysis, and computational fluid dynamics modeling have all demonstrated the viability of the GTX concept, yet significant technical issues and challenges still remain. Our research effort develops a unique capability for dynamic CFD simulation of complete high-speed propulsion devices and focuses this technology toward analysis of the GTX response during critical mode transition events. Our principal attention is focused on Mode 1/Mode 2 operation, in which initial rocket propulsion is transitioned into thermal-throat ramjet propulsion. A critical element of the GTX concept is the use of an Independent Ramjet Stream (IRS) cycle to provide propulsion at Mach numbers less than 3. In the IRS cycle, rocket thrust is initially used for primary power, and the hot rocket plume is used as a flame-holding mechanism for hydrogen fuel injected into the secondary air stream. A critical aspect is the establishment of a thermal throat in the secondary stream through the combination of area reduction effects and combustion-induced heat release. This is a necessity to enable the power-down of the rocket and the eventual shift to ramjet mode. Our focus in this first year of the grant has been in three areas, each progressing directly toward the key initial goal of simulating thermal throat formation during the IRS cycle: CFD algorithm development; simulation of Mode 1 experiments conducted at Glenn's Rig 1 facility; and IRS cycle simulations. The remainder of this report discusses each of these efforts in detail and presents a plan of work for the next year.

Rocket-Based Combined Cycle Flowpath Testing for Modes 1 and 4

NASA Technical Reports Server (NTRS)

Rice, Tharen

2002-01-01

Under sponsorship of the NASA Glenn Research Center (NASA GRC), the Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and built a five-inch diameter, Rocket-Based Combined Cycle (RBCC) engine to investigate mode 1 and mode 4 engine performance as well as Mach 4 inlet performance. This engine was designed so that engine area and length ratios were similar to the NASA GRC GTX engine is shown. Unlike the GTX semi-circular engine design, the APL engine is completely axisymmetric. For this design, a traditional rocket thruster was installed inside of the scramjet flowpath, along the engine centerline. A three part test series was conducted to determine Mode I and Mode 4 engine performance. In part one, testing of the rocket thruster alone was accomplished and its performance determined (average Isp efficiency = 90%). In part two, Mode 1 (air-augmented rocket) testing was conducted at a nominal chamber pressure-to-ambient pressure ratio of 100 with the engine inlet fully open. Results showed that there was neither a thrust increment nor decrement over rocket-only thrust during Mode 1 operation. In part three, Mode 4 testing was conducted with chamber pressure-to-ambient pressure ratios lower than desired (80 instead of 600) with the inlet fully closed. Results for this testing showed a performance decrease of 20% as compared to the rocket-only testing. It is felt that these results are directly related to the low pressure ratio tested and not the engine design. During this program, Mach 4 inlet testing was also conducted. For these tests, a moveable centerbody was tested to determine the maximum contraction ratio for the engine design. The experimental results agreed with CFD results conducted by NASA GRC, showing a maximum geometric contraction ratio of approximately 10.5. This report details the hardware design, test setup, experimental results and data analysis associated with the aforementioned tests.

Rocket-Based Combined Cycle Engine Technology Development: Inlet CFD Validation and Application

NASA Technical Reports Server (NTRS)

DeBonis, J. R.; Yungster, S.

1996-01-01

A CFD methodology has been developed for inlet analyses of Rocket-Based Combined Cycle (RBCC) Engines. A full Navier-Stokes analysis code, NPARC, was used in conjunction with pre- and post-processing tools to obtain a complete description of the flow field and integrated inlet performance. This methodology was developed and validated using results from a subscale test of the inlet to a RBCC 'Strut-Jet' engine performed in the NASA Lewis 1 x 1 ft. supersonic wind tunnel. Results obtained from this study include analyses at flight Mach numbers of 5 and 6 for super-critical operating conditions. These results showed excellent agreement with experimental data. The analysis tools were also used to obtain pre-test performance and operability predictions for the RBCC demonstrator engine planned for testing in the NASA Lewis Hypersonic Test Facility. This analysis calculated the baseline fuel-off internal force of the engine which is needed to determine the net thrust with fuel on.

Research Technology

NASA Image and Video Library

2004-04-15

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

Thermal analysis of regenerative-cooled pylon in multi-mode rocket based combined cycle engine

NASA Astrophysics Data System (ADS)

Yan, Dekun; He, Guoqiang; Li, Wenqiang; Zhang, Duo; Qin, Fei

2018-07-01

Combining pylon injector with rocket is an effective method to achieve efficient mixing and combustion in the RBCC engine. This study designs a fuel pylon with active cooling structure, and numerically investigates the coupled heat transfer between active cooling process in the pylon and combustion in the combustor in different modes. Effect of the chemical reaction of the fuel on the flow, heat transfer and physical characteristics is also discussed. The numerical results present a good agreement with the experimental data. Results indicate that drastic supplementary combustion caused by rocket gas and secondary combustion caused by the fuel injection from the pylon result in severe thermal load on the pylon. Although regenerative cooling without cracking can reduce pylon's temperature below the allowable limit, a high-temperature area appears in the middle and nail section of the pylon due to the coolant's insufficient convective heat transfer coefficient. Comparatively, endothermic cracking can provide extra chemical heat sink for the coolant and low velocity contributes to prolong the reaction time to increase the heat absorption from chemical reaction, which further lowers and unifies the pylon surface temperature.

Design Rules and Issues with Respect to Rocket Based Combined Cycles

DTIC Science & Technology

2010-09-01

cause thrust augmentation due to the ejector effects, which in turn, can reduce the requirement for the rocket engine output. In the speed regime with...should produce sufficient thrust to takeoff and to overcome the drag at transonic regime. When embedded into a flow pass, the rocket exhaust can...between the ejector -jet operation and ramjet operation, between the ramjet operations at various flight conditions, and between the ramjet operation and

The Trailblazer Program

NASA Technical Reports Server (NTRS)

Trefney, Charles J.

1999-01-01

This paper presents the "Three Pillars of Success" for the Trailblazer Program. The topics include: 1) The "Rocket Equation" for SSTO (Single Stage To Orbit); 2) The Rocket I* Barrier; 3) Rocket-Based Combined-Cycle Engine; 4) Potential for Reusability; 5) Factors Mitigating RBCC Performance; 6) The "Trailblazer" Program; 7) Trailblazer Performance Goals; 8) Trailblazer Reference Vehicle; and 9) Trailblazer Program Architecture.

Experimental Studies of the Heat Transfer to RBCC Rocket Nozzles for CFD Application to Design Methodologies

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

1999-01-01

Rocket thrusters for Rocket Based Combined Cycle (RBCC) engines typically operate with hydrogen/oxygen propellants in a very compact space. Packaging considerations lead to designs with either axisymmetric or two-dimensional throat sections. Nozzles tend to be either two- or three-dimensional. Heat transfer characteristics, particularly in the throat, where the peak heat flux occurs, are not well understood. Heat transfer predictions for these small thrusters have been made with one-dimensional analysis such as the Bartz equation or scaling of test data from much larger thrusters. The current work addresses this issue with an experimental program that examines the heat transfer characteristics of a gaseous oxygen (GO2)/gaseous hydrogen (GH2) two-dimensional compact rocket thruster. The experiments involved measuring the axial wall temperature profile in the nozzle region of a water-cooled gaseous oxygen/gaseous hydrogen rocket thruster at a pressure of 3.45 MPa. The wall temperature measurements in the thruster nozzle in concert with Bartz's correlation are utilized in a one-dimensional model to obtain axial profiles of nozzle wall heat flux.

The Pulse Detonation Rocket Induced MHD Ejector (PDRIME) Concept (Preprint)

DTIC Science & Technology

2008-06-10

flight applications. Thrust augmentation , such as PDE- ejector configurations, can potentially alleviate this problem. Here, we study the potential...flow, to assist in augmentation of the thrust . Ejectors typically transfer energy between streams through shear stress between separate flow streams...and the ejector operates. This is one of several configurations in which the PDRIME concept could be used for thrust augmentation in advanced

Operationally efficient propulsion system study (OEPSS) data book. Volume 10; Air Augmented Rocket Afterburning

NASA Technical Reports Server (NTRS)

Farhangi, Shahram; Trent, Donnie (Editor)

1992-01-01

A study was directed towards assessing viability and effectiveness of an air augmented ejector/rocket. Successful thrust augmentation could potentially reduce a multi-stage vehicle to a single stage-to-orbit vehicle (SSTO) and, thereby, eliminate the associated ground support facility infrastructure and ground processing required by the eliminated stage. The results of this preliminary study indicate that an air augmented ejector/rocket propulsion system is viable. However, uncertainties resulting from simplified approach and assumptions must be resolved by further investigations.

Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

NASA Technical Reports Server (NTRS)

Fernandez, Rene; Trefny, Charles J.; Thomas, Scott R.; Bulman, Mel J.

2001-01-01

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

Mach 5 to 7 RBCC Propulsion System Testing at NASA-LeRC HTF

NASA Technical Reports Server (NTRS)

Perkins, H. Douglas; Thomas, Scott R.; Pack, William D.

1996-01-01

A series of Mach 5 to 7 freejet tests of a Rocket Based Combined Cycle (RBCC) engine were cnducted at the NASA Lewis Research Center (LERC) Hypersonic Tunnel Facility (HTF). This paper describes the configuration and operation of the HTF and the RBCC engine during these tests. A number of facility support systems are described which were added or modified to enhance the HTF test capability for conducting this experiment. The unfueled aerodynamic perfor- mance of the RBCC engine flowpath is also presented and compared to sub-scale test results previously obtained in the NASA LERC I x I Supersonic Wind Tunnel (SWT) and to Computational Fluid Dynamic (CFD) analysis results. This test program demonstrated a successful configuration of the HTF for facility starting and operation with a generic RBCC type engine and an increased range of facility operating conditions. The ability of sub-scale testing and CFD analysis to predict flowpath performance was also shown. The HTF is a freejet, blowdown propulsion test facility that can simulate up to Mach 7 flight conditions with true air composition. Mach 5, 6, and 7 facility nozzles are available, each with an exit diameter of 42 in. This combination of clean air, large scale, and Mach 7 capabilities is unique to the HTF. This RBCC engine study is the first engine test program conducted at the HTF since 1974.

Aero-Thermo-Structural Analysis of Inlet for Rocket Based Combined Cycle Engines

NASA Technical Reports Server (NTRS)

Shivakumar, K. N.; Challa, Preeti; Sree, Dave; Reddy, Dhanireddy R. (Technical Monitor)

2000-01-01

NASA has been developing advanced space transportation concepts and technologies to make access to space less costly. One such concept is the reusable vehicles with short turn-around times. The NASA Glenn Research Center's concept vehicle is the Trailblazer powered by a rocket-based combined cycle (RBCC) engine. Inlet is one of the most important components of the RBCC engine. This paper presents fluid flow, thermal, and structural analysis of the inlet for Mach 6 free stream velocity for fully supersonic and supercritical with backpressure conditions. The results concluded that the fully supersonic condition was the most severe case and the largest stresses occur in the ceramic matrix composite layer of the inlet cowl. The maximum tensile and the compressive stresses were at least 3.8 and 3.4, respectively, times less than the associated material strength.

Ignition and flame stabilization of a strut-jet RBCC combustor with small rocket exhaust.

PubMed

Hu, Jichao; Chang, Juntao; Bao, Wen

2014-01-01

A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes.

Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

PubMed Central

2014-01-01

A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 2; Unsteady Analyses and Risk Assessment

NASA Technical Reports Server (NTRS)

Ahuja, Vineet; Hosangadi, Ashvin; Allgood, Daniel

2008-01-01

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Design Support of the feasibility of operating conditions and procedures is critical in such cases due to the possibility of startup/shutdown transients, moving shock structures, unsteady shock-boundary layer interactions and engine and diffuser unstart modes that can result in catastrophic failure. Analyses of such systems is difficult due to resolution requirements needed to accurately capture moving shock structures, shock-boundary layer interactions, two-phase flow regimes and engine unstart modes. In a companion paper, we will demonstrate with the use of CFD, steady analyses advanced capability to evaluate supersonic diffuser and steam ejector performance in the sub-scale A-3 facility. In this paper we will address transient issues with the operation of the facility especially at startup and shutdown, and assess risks related to afterburning due to the interaction of a fuel rich plume with oxygen that is a by-product of the steam ejectors. The primary areas that will be addressed in this paper are: (1) analyses of unstart modes due to flow transients especially during startup/ignition, (2) engine safety during the shutdown process (3) interaction of steam ejectors with the primary plume i.e. flow transients as well as probability of afterburning. In this abstract we discuss unsteady analyses of the engine shutdown process. However, the final paper will include analyses of a staged startup, drawdown of the engine test cell pressure, and risk assessment of potential afterburning in the facility. Unsteady simulations have been carried out to study the engine shutdown process in the facility and understand the physics behind the interactions between the steam ejectors, the test cell and the supersonic diffuser. As a first approximation, to understand the dominant unsteady mechanisms in the engine test cell and the supersonic diffuser, the turning duct in the facility was removed. As the engine loses power a rarefaction wave travels downstream that disrupts the shock cell structure in the supersonic diffuser. Flow from the test cell is seen to expand into the supersonic diffuser section and re-pressurizes the area around the nozzle along with a upstream traveling compression wave that emanates from near the first stage ejectors. Flow from the first stage ejector expands to the center of the duct and a new shock train is formed between the first and second stage ejectors. Both stage ejectors keep the facility pressurized and prevent any large amplitude pressure fluctuations from affecting the engine nozzle. The resultant pressure loads the nozzle experiences in the shutdown process are small.

Research Technology

NASA Image and Video Library

2004-04-15

Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

Experimental Investigation of an Integrated Strut-Rocket/Scramjet Operating at Mach 4.0 and 6.5 Conditions

NASA Technical Reports Server (NTRS)

Hawk, Clark; Nelson, Karl

1998-01-01

A series of tests were conducted to investigate RBCC performance at ramjet and scramjet conditions. The hardware consisted of a linear strut-rocket manufactured by Aerojet and a dual-mods scramjet combustor. The hardware was tested at NASA Langley Research Center in the Direct Connect Supersonic Combustion Test Facility at Mach 4.0 and 6.5 simulated flight conditions.

Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

NASA Astrophysics Data System (ADS)

Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

2016-10-01

This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

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

NASA Technical Reports Server (NTRS)

Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

1959-01-01

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.

Computational Fluid Dynamic Modeling of Rocket Based Combined Cycle Engine Flowfields

NASA Technical Reports Server (NTRS)

Daines, Russell L.; Merkle, Charles L.

1994-01-01

Computational Fluid Dynamic techniques are used to study the flowfield of a fixed geometry Rocket Based Combined Cycle engine operating in rocket ejector mode. Heat addition resulting from the combustion of injected fuel causes the subsonic engine flow to choke and go supersonic in the slightly divergent combustor-mixer section. Reacting flow computations are undertaken to predict the characteristics of solutions where the heat addition is determined by the flowfield. Here, adaptive gridding is used to improve resolution in the shear layers. Results show that the sonic speed is reached in the unheated portions of the flow first, while the heated portions become supersonic later. Comparison with results from another code show reasonable agreement. The coupled solutions show that the character of the combustion-based thermal choking phenomenon can be controlled reasonably well such that there is opportunity to optimize the length and expansion ratio of the combustor-mixer.

Rocket-Based Combined Cycle Activities in the Advanced Space Transportation Program Office

NASA Technical Reports Server (NTRS)

Hueter, Uwe; Turner, James

1999-01-01

NASA's Office of Aero-Space Technology (OAST) has established three major goals, referred to as, "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center (MSFC) in Huntsville, Ala. focuses on future space transportation technologies Under the "Access to Space" pillar. The Core Technologies Project, part of ASTP, focuses on the reusable technologies beyond those being pursued by X-33. One of the main activities over the past two and a half years has been on advancing the rocket-based combined cycle (RBCC) technologies. In June of last year, activities for reusable launch vehicle (RLV) airframe and propulsion technologies were initiated. These activities focus primarily on those technologies that support the decision to determine the path this country will take for Space Shuttle and RLV. This year, additional technology efforts in the reusable technologies will be awarded. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion.

Development of a numerical tool to study the mixing phenomenon occurring during mode one operation of a multi-mode ejector-augmented pulsed detonation rocket engine

NASA Astrophysics Data System (ADS)

Dawson, Joshua

A novel multi-mode implementation of a pulsed detonation engine, put forth by Wilson et al., consists of four modes; each specifically designed to capitalize on flow features unique to the various flow regimes. This design enables the propulsion system to generate thrust through the entire flow regime. The Multi-Mode Ejector-Augmented Pulsed Detonation Rocket Engine operates in mode one during take-off conditions through the acceleration to supersonic speeds. Once the mixing chamber internal flow exceeds supersonic speed, the propulsion system transitions to mode two. While operating in mode two, supersonic air is compressed in the mixing chamber by an upstream propagating detonation wave and then exhausted through the convergent-divergent nozzle. Once the velocity of the air flow within the mixing chamber exceeds the Chapman-Jouguet Mach number, the upstream propagating detonation wave no longer has sufficient energy to propagate upstream and consequently the propulsive system shifts to mode three. As a result of the inability of the detonation wave to propagate upstream, a steady oblique shock system is established just upstream of the convergent-divergent nozzle to initiate combustion. And finally, the propulsion system progresses on to mode four operation, consisting purely of a pulsed detonation rocket for high Mach number flight and use in the upper atmosphere as is needed for orbital insertion. Modes three and four appear to be a fairly significant challenge to implement, while the challenge of implementing modes one and two may prove to be a more practical goal in the near future. A vast number of potential applications exist for a propulsion system that would utilize modes one and two, namely a high Mach number hypersonic cruise vehicle. There is particular interest in the dynamics of mode one operation, which is the subject of this research paper. Several advantages can be obtained by use of this technology. Geometrically the propulsion system is fairly simple and as a result of the rapid combustion process the engine cycle is more efficient compared to its combined cycle counterparts. The flow path geometry consists of an inlet system, followed just downstream by a mixing chamber where an ejector structure is placed within the flow path. Downstream of the ejector structure is a duct leading to a convergent-divergent nozzle. During mode one operation and within the ejector, products from the detonation of a stoichiometric hydrogen/air mixture are exhausted directly into the surrounding secondary air stream. Mixing then occurs between both the primary and secondary flow streams, at which point the air mass containing the high pressure, high temperature reaction products is convected downstream towards the nozzle. The engine cycle is engineered to a specific number of detonations per second, creating the pulsating characteristic of the primary flow. The pulsing nature of the primary flow serves as a momentum augmentation, enhancing the thrust and specific impulse at low speeds. Consequently it is necessary to understand the transient mixing process between the primary and secondary flow streams occurring during mode one operation. Using OPENFOAMRTM, an analytic tool is developed to simulate the dynamics of the turbulent detonation process along with detailed chemistry in order to understand the physics involved with the stream interactions. The computational code has been developed within the framework of OPENFOAMRTM, an open-source alternative to commercial CFD software. A conservative formulation of the Farve averaged Navier-Stokes equations is implemented to facilitate programming and numerical stability. Time discretization is accomplished by using the Crank-Nicolson method, achieving second order convergence in time. Species mass fraction transport equations are implemented and a Seulex ODE solver was used to resolve the system of ordinary differential equations describing the hydrogen-air reaction mechanism detailed in Appendix A. The Seulex ODE solution algorithm is an extrapolation method based on the linearly implicit Euler method with step size control. A second order total variation diminishing method with a modified Sweby flux limiter was used for space discretization. And finally the use of operator splitting (PISO algorithm, and chemical kinetics) is essential due to the significant differences in characteristic time scales evolving simultaneously in turbulent reactive flow. Capturing the turbulent nature of the combustion process was done using the k-o-SST turbulence model, as formulated by Mentor [1]. Mentor's formulation is well suited to resolve the boundary layer while remaining relatively insensitive to freestream conditions, blending the merits of both the k-o and k-epsilon models. Further development of the tool is possible, most notably with the Numerical Propulsion System Simulation application. NPSS allows the user to take advantage of a "zooming" functionality in which high fidelity models of engine components can be integrated into NPSS models, allowing for a more robust propulsion system simulation.

Investigation of an ejector heat pump by analytical methods

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hsu, C.T.

1984-07-01

Using existing theories of ejector design, the optimum geometry of a high-efficiency ejector - including mixing section cross-sectional area, mass flow entrainment rate, ejector efficiency, and overall COP - for a heat pump cycle was determined. A parametric study was performed to evaluate the COP values for different operating conditions. A sensitivity study determined th effects of nozzle efficiency and diffuser efficiency on the overall ejector heat pump COP. The off-design study estimated the COP for an ejector heat pump operating at off-design conditions. Refrigerants 11, 113, and 114 are three of the halocarbons which best satisfy the criteria formore » an ejector heat pump system. The estimated COPs were 0.3 for the cooling mode and 1.3 for the heating mode at standard operating conditions: a boiler temperature of 93.3/sup 0/C (200/sup 0/F), a condenser temperature of 43.3/sup 0/C (110/sup 0/F), and an evaporator temperature of 10/sup 0/C (50/sup 0/F). Based on the same operating conditions, an optimum ejector geometry was estimated for each of the refrigerants R-11 and R-113. Since the COP values for heating obtained in this analysis are greater than unity, the performance of an ejector heat pump operating in the heating mode should be competitive with that of oil- or gas-fired furnaces or electrical resistance heaters.« less

Development Activities on Airbreathing Combined Cycle Engines

NASA Technical Reports Server (NTRS)

McArthur, J. Craig; Lyles, Garry (Technical Monitor)

2000-01-01

Contents include the following: Advanced reusable transportation(ART); aerojet and rocketdyne tests, RBCC focused concept flowpaths,fabricate flight weigh, test select components, document ART project, Istar (Integrated system test of an airbreathing rocket); combined cycle propulsion testbed;hydrocarbon demonstrator tracebility; Istar engine system and vehicle system closure study; and Istar project planning.

Space Transportation Technology Workshop: Propulsion Research and Technology

NASA Technical Reports Server (NTRS)

2000-01-01

This viewgraph presentation gives an overview of the Space Transportation Technology Workshop topics, including Propulsion Research and Technology (PR&T) project level organization, FY 2001 - 2006 project roadmap, points of contact, foundation technologies, auxiliary propulsion technology, PR&T Low Cost Turbo Rocket, and PR&T advanced reusable technologies RBCC test bed.

Subsonic Performance of Ejector Systems

NASA Astrophysics Data System (ADS)

Weil, Samuel

Combined cycle engines combining scramjets with turbo jets or rockets can provide efficient hypersonic flight. Ejectors have the potential to increase the thrust and efficiency of combined cycle engines near static conditions. A computer code was developed to support the design of a small-scale, turbine-based combined cycle demonstrator with an ejector, built around a commercially available turbojet engine. This code was used to analyze the performance of an ejector system built around a micro-turbojet. With the use of a simple ejector, net thrust increases as large as 20% over the base engine were predicted. Additionally the specific fuel consumption was lowered by 10%. Increasing the secondary to primary area ratio of the ejector lead to significant improvements in static thrust, specific fuel consumption (SFC), and propulsive efficiency. Further ejector performance improvements can be achieved by using a diffuser. Ejector performance drops off rapidly with increasing Mach number. The ejector has lower thrust and higher SFC than the turbojet core at Mach numbers above 0.2. When the nozzle chokes a significant drop in ejector performance is seen. When a diffuser is used, higher Mach numbers lead to choking in the mixer and a shock in the nozzle causing a significant decrease in ejector performance. Evaluation of different turbo jets shows that ejector performance depends significantly on the properties of the turbojet. Static thrust and SFC improvements can be achieved with increasing ejector area for all engines, but size of increase and change in performance at higher Mach numbers depend heavily on the turbojet. The use of an ejector in a turbine based combined cycle configuration also increases performance at static conditions with a thrust increase of 5% and SFC decrease of 5% for the tested configuration.

High-Temperature Polymer Composites Tested for Hypersonic Rocket Combustor Backup Structure

NASA Technical Reports Server (NTRS)

Sutter, James K.; Shin, E. Eugene; Thesken, John C.; Fink, Jeffrey E.

2005-01-01

Significant component weight reductions are required to achieve the aggressive thrust-toweight goals for the Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-toorbit vehicle. A collaboration between the NASA Glenn Research Center and Boeing Rocketdyne was formed under the Higher Operating Temperature Propulsion Components (HOTPC) program and, currently, the Ultra-Efficient Engine Technology (UEET) Project to develop carbon-fiber-reinforced high-temperature polymer matrix composites (HTPMCs). This program focused primarily on the combustor backup structure to replace all metallic support components with a much lighter polymer-matrixcomposite- (PMC-) titanium honeycomb sandwich structure.

I(sup STAR), NASA's Next Step in Air-Breathing Propulsion for Space Access

NASA Technical Reports Server (NTRS)

Hutt, John J.; McArthur, Craig; Cook, Stephen (Technical Monitor)

2001-01-01

The United States' National Aeronautics and Space Administration (NASA) has established a strategic plan for future activities in space. A primary goal of this plan is to make drastic improvements in the cost and safety of earth to low-earth-orbit transportation. One approach to achieving this goal is through the development of highly reusable, highly reliable space transportation systems analogous to the commercial airline system. In the year 2000, NASA selected the Rocket Based Combined Cycle (RBCC) engine as the next logical step towards this goal. NASA will develop a complete flight-weight, pump-fed engine system under the Integrated System Test of an Airbreathing Rocket (I(sup STAR)) Project. The objective of this project is develop a reusable engine capable of self-powering a vehicle through the air-augmented rocket, ramjet and scramjet modes required in all RBCC based operational vehicle concepts. The project is currently approved and funded to develop the engine through ground test demonstration. Plans are in place to proceed with flight demonstration pending funding approval. The project is in formulation phase and the Preliminary Requirements Review has been completed. The engine system and vehicle have been selected at the conceptual level. The I(sup STAR) engine concept is based on an air-breathing flowpath downselected from three configurations evaluated in NASA's Advanced Reusable Technology contract. The selected flowpath features rocket thrust chambers integrated into struts separating modular flowpath ducts, a variable geometry inlet, and a thermally choked throat. The engine will be approximately 220 inches long and 79 inches wide and fueled with a hydrocarbon fuel using liquid oxygen as the primary oxidizer candidate. The primary concept for the pump turbine drive is pressure-fed catalyzed hydrogen peroxide. In order to control costs, the flight demonstration vehicle will be launched from a B-52 aircraft. The vehicle concept is based on the Air Breathing Launch Vehicle 4 (ABLV4) lifting body configuration which has design heritage from NASA's NASP Program. The vehicle will be designed to accelerate from Mach 0.8 to Mach 7 and will be equipped with landing gear for horizontal landing. The complete vehicle, including the engine, will be designed for 25 flights and will be approximately 33 feet long with a total vehicle weight of approximately 25000 lbs.

Generation 1.5 High Speed Civil Transport (HSCT) Exhaust Nozzle Program

NASA Technical Reports Server (NTRS)

Thayer, E. B.; Gamble, E. J.; Guthrie, A. R.; Kehret, D. F.; Barber, T. J.; Hendricks, G. J.; Nagaraja, K. S.; Minardi, J. E.

2004-01-01

The objective of this program was to conduct an experimental and analytical evaluation of low noise exhaust nozzles suitable for future High-Speed Civil Transport (HSCT) aircraft. The experimental portion of the program involved parametric subscale performance model tests of mixer/ejector nozzles in the takeoff mode, and high-speed tests of mixer/ejectors converted to two-dimensional convergent-divergent (2-D/C-D), plug, and single expansion ramp nozzles (SERN) in the cruise mode. Mixer/ejector results show measured static thrust coefficients at secondary flow entrainment levels of 70 percent of primary flow. Results of the high-speed performance tests showed that relatively long, straight-wall, C-D nozzles could meet supersonic cruise thrust coefficient goal of 0.982; but the plug, ramp, and shorter C-D nozzles required isentropic contours to reach the same level of performance. The computational fluid dynamic (CFD) study accurately predicted mixer/ejector pressure distributions and shock locations. Heat transfer studies showed that a combination of insulation and convective cooling was more effective than film cooling for nonafterburning, low-noise nozzles. The thrust augmentation study indicated potential benefits for use of ejector nozzles in the subsonic cruise mode if the ejector inlet contains a sonic throat plane.

Operating and hydrodynamic characteristics of a reversed flow jet loop bioreactor (RFJLB) with ejector.

PubMed

Wagh, Sameer M; Koranne, Kishore V; Sonolikar, Ram L

2012-04-01

The hydrodynamic characteristics of RFJLB was studied with superficial liquid velocity (Ul), nozzle diameter (Dn) and nozzle height (Hn) in the range of 0.0293-0.094m/s, 17.4-22.0mm and 50-400mm, respectively. For Dn=17.4mm, Hn=50 and 200mm, with ejector mode and regular operating procedure i.e. simultaneous entry of gas with increasing liquid velocity, had limitation of not establishing the circulation loop. To overcome this limitation a modified operating procedure i.e. entry of gas after established liquid circulation loop is proposed. Also the comparison of gas holdups with ejector and injector mode proves the effectiveness of ejector mode and can eliminate the supply of compressed gas. Thus proper choice of Dn, Hn and also the operating procedure becomes necessary. Copyright © 2012 Elsevier Ltd. All rights reserved.

NASA's Advanced Propulsion Technology Activities for Third Generation Fully Reusable Launch Vehicle Applications

NASA Technical Reports Server (NTRS)

Hueter, Uwe

2000-01-01

NASA's Office of Aeronautics and Space Transportation Technology (OASTT) established the following three major goals, referred to as "The Three Pillars for Success": Global Civil Aviation, Revolutionary Technology Leaps, and Access to Space. The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville, Ala. focuses on future space transportation technologies under the "Access to Space" pillar. The Propulsion Projects within ASTP under the investment area of Spaceliner100, focus on the earth-to-orbit (ETO) third generation reusable launch vehicle technologies. The goals of Spaceliner 100 is to reduce cost by a factor of 100 and improve safety by a factor of 10,000 over current conditions. The ETO Propulsion Projects in ASTP, are actively developing combination/combined-cycle propulsion technologies that utilized airbreathing propulsion during a major portion of the trajectory. System integration, components, materials and advanced rocket technologies are also being pursued. Over the last several years, one of the main thrusts has been to develop rocket-based combined cycle (RBCC) technologies. The focus has been on conducting ground tests of several engine designs to establish the RBCC flowpaths performance. Flowpath testing of three different RBCC engine designs is progressing. Additionally, vehicle system studies are being conducted to assess potential operational space access vehicles utilizing combined-cycle propulsion systems. The design, manufacturing, and ground testing of a scale flight-type engine are planned. The first flight demonstration of an airbreathing combined cycle propulsion system is envisioned around 2005. The paper will describe the advanced propulsion technologies that are being being developed under the ETO activities in the ASTP program. Progress, findings, and future activities for the propulsion technologies will be discussed.

A-3 Test Stand construction moves forward

NASA Image and Video Library

2010-07-13

Work on the A-3 Test Stand at Stennis Space Center took a step forward in July with delivery of the first-stage steam ejector July 13. Stennis employees are shown preparing the ejector to be lifted into place on the test stand. When activated in 2012, the A-3 Test Stand will allow operators to test rocket engines at simulated altitudes of 100,000 feet, a critical feature for next-generation engines that will take humans beyond low-Earth orbit once more.

CPU and GPU-based Numerical Simulations of Combustion Processes

DTIC Science & Technology

2012-04-27

Distribution unlimited UCLA MAE Research and Technology Review April 27, 2012 Magnetohydrodynamic Augmentation of the Pulse Detonation Rocket Engines...Pulse Detonation Rocket-Induced MHD Ejector (PDRIME) – Energy extract from exhaust flow by MHD generator – Seeded air stream acceleration by MHD...accelerator for thrust enhancement and control • Alternative concept: Magnetic piston – During PDE blowdown process, MHD extracts energy and

Qualification Test of the Thiokol TE-M-364-19 Solid-Propellant Rocket Motor (S/N 19006)

DTIC Science & Technology

1977-05-01

cell by a steam ejector operating in series with the ETF exhaust gas compressors. During the motor firing, the motor exhaust gases were used as a...driving gas for the 42-in.-diam, water-cooled, ejector-diffuser system incorporating a 24-deg (half-angle) conical inlet to maintain test cell pressure...after Ignition, sec 0.5 0.6 0.7 Figure 4. Variation of thrust and chamber pressure during motor ignition. - CO Q_ OH LU CO TL cr x CJ 1400

Research on shock wave characteristics in the isolator of central strut rocket-based combined cycle engine under Ma5.5

NASA Astrophysics Data System (ADS)

Wei, Xianggeng; Xue, Rui; Qin, Fei; Hu, Chunbo; He, Guoqiang

2017-11-01

A numerical calculation of shock wave characteristics in the isolator of central strut rocket-based combined cycle (RBCC) engine fueled by kerosene was carried out in this paper. A 3D numerical model was established by the DES method. The kerosene chemical kinetic model used the 9-component and 12-step simplified mechanism model. Effects of fuel equivalence ratio, inflow total temperature and central strut rocket on-off on shock wave characteristics were studied under Ma5.5. Results demonstrated that with the increase of equivalence ratio, the leading shock wave moves toward upstream, accompanied with higher possibility of the inlet unstart. However, the leading shock wave moves toward downstream as the inflow total temperature rises. After the central strut rocket is closed, the leading shock wave moves toward downstream, which can reduce risks of the inlet unstart. State of the shear layer formed by the strut rocket jet flow and inflow can influence the shock train structure significantly.

An evaluation of composite propulsion for single-stage-to-orbit vehicles designed for horizontal take-off

NASA Technical Reports Server (NTRS)

Martin, J. A.

1977-01-01

Composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off. Trajectory, geometric, and mass analyses were performed to establish the orbital payload capability of six engines. The results indicated that none of the engines performed adequately to deliver payloads to orbit as analyzed. The single-stage turbine and oxidizer-rich gas generator resulted in a low engine specific impulse, and the performance increment of the ejector subsystem was less than that of a separate rocket system with a high combustion pressure. There was a benefit from incorporating a fan into the engine, and removal of the fan from the airstream during the ramjet mode increased the orbital payload capability.

Air-Breathing Launch Vehicle Technology Being Developed

NASA Technical Reports Server (NTRS)

Trefny, Charles J.

2003-01-01

Of the technical factors that would contribute to lowering the cost of space access, reusability has high potential. The primary objective of the GTX program is to determine whether or not air-breathing propulsion can enable reusable single-stage-to-orbit (SSTO) operations. The approach is based on maturation of a reference vehicle design with focus on the integration and flight-weight construction of its air-breathing rocket-based combined-cycle (RBCC) propulsion system.

Multidisciplinary Design Techniques Applied to Conceptual Aerospace Vehicle Design. Ph.D. Thesis Final Technical Report

NASA Technical Reports Server (NTRS)

Olds, John Robert; Walberg, Gerald D.

1993-01-01

Multidisciplinary design optimization (MDO) is an emerging discipline within aerospace engineering. Its goal is to bring structure and efficiency to the complex design process associated with advanced aerospace launch vehicles. Aerospace vehicles generally require input from a variety of traditional aerospace disciplines - aerodynamics, structures, performance, etc. As such, traditional optimization methods cannot always be applied. Several multidisciplinary techniques and methods were proposed as potentially applicable to this class of design problem. Among the candidate options are calculus-based (or gradient-based) optimization schemes and parametric schemes based on design of experiments theory. A brief overview of several applicable multidisciplinary design optimization methods is included. Methods from the calculus-based class and the parametric class are reviewed, but the research application reported focuses on methods from the parametric class. A vehicle of current interest was chosen as a test application for this research. The rocket-based combined-cycle (RBCC) single-stage-to-orbit (SSTO) launch vehicle combines elements of rocket and airbreathing propulsion in an attempt to produce an attractive option for launching medium sized payloads into low earth orbit. The RBCC SSTO presents a particularly difficult problem for traditional one-variable-at-a-time optimization methods because of the lack of an adequate experience base and the highly coupled nature of the design variables. MDO, however, with it's structured approach to design, is well suited to this problem. The result of the application of Taguchi methods, central composite designs, and response surface methods to the design optimization of the RBCC SSTO are presented. Attention is given to the aspect of Taguchi methods that attempts to locate a 'robust' design - that is, a design that is least sensitive to uncontrollable influences on the design. Near-optimum minimum dry weight solutions are determined for the vehicle. A summary and evaluation of the various parametric MDO methods employed in the research are included. Recommendations for additional research are provided.

Steam jet ejectors for the process industries. [Glossary included

DOE Office of Scientific and Technical Information (OSTI.GOV)

Power, R.B.

1994-01-01

Steam jet ejectors were for many years the workhorse of the chemical process industries for producing vacuum. With increasing emphasis on stricter pollution control, their use was curtailed. There are still many applications, however, such as those with large capacity requirements, where ejectors are the only equipment that can produce sufficient vacuum. Chapter 1 is a short overview on how to use the text. Chapter 2 discusses what an ejector is and how it works. How ejector stages work is reviewed in Chapter 3. Engineering calculations for ejector stages is thoroughly discussed in Chapter 4. In Chapter 5, contact andmore » surface condensers are reviewed, and calculation procedures are presented. The various types of pressure control are discussed in Chapter 6. Chapter 7 is an excellent review of installation of ejector vacuum systems. The final chapter of Part 2 (Chapters 3--8) thoroughly covers all aspects of operation, testing, troubleshooting and maintenance. Part 3, consisting of two chapters, is devoted to specifying and purchasing steam jet ejectors. Part 4 on other ejector applications and upgrading ejector usage also consists of two chapters. Chapter 11 reviews steam-jet refrigeration, steam-jet and gas-jet compressors, liquid jet eductors, desuperheaters, special design situations, and designing one's own systems. Upgrading of existing ejector procedures and hardware is reviewed in Chapter 12. The 12 appendixes cover: physical properties of common fluids; handy vacuum engineering data and rules of thumb; SI unit conversions; sizing air and steam metering orifices for testing; drill sizes; ejector operating costs and design optimization; forms for ejector calculations, tests, and inspections; instructions for preparing ejector specifications; test kit contents list; ejector manufacturers and suppliers of referenced hardware and information; and failure modes and symptoms.« less

Validation of the NCC Code for Staged Transverse Injection and Computations for a RBCC Combustor

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Liu, Nan-Suey

2005-01-01

The NCC code was validated for a case involving staged transverse injection into Mach 2 flow behind a rearward facing step. Comparisons with experimental data and with solutions from the FPVortex code was then used to perform computations to study fuel-air mixing for the combustor of a candidate rocket based combined cycle engine geometry. Comparisons with a one-dimensional analysis and a three-dimensional code (VULCAN) were performed to assess the qualitative and quantitative performance of the NCC solver.

Deep Throttle Turbopump Technology Testing

NASA Technical Reports Server (NTRS)

Ferguson, T. V.; Guinzburg, A.; McGlynn, R. D.; Williams, M.

2002-01-01

The objectives of this viewgraph presentation were to: (1) enhance and demonstrate critical technologies in support of planned RBCC flight test programs; and (2) obtain knowledge of wide flow range as it is applicable to liquid rocket engine turbopumps operating over extreme throttle ranges. This program was set up to demonstrate wide flow range diffuser technologies. The testing phase of the contract to provide data to anchor initial designs was partially successful. Data collected suggest flow phenomena exists at off-design flow rates.

An analytical study of hybrid ejector/internal combustion engine-driven heat pumps

DOE Office of Scientific and Technical Information (OSTI.GOV)

Murphy, R.W.

1988-01-01

Because ejectors can combine high reliability with low maintenance cost in a package requiring little capital investment, they may provide attractive heat pumping capability in situations where the importance of their inefficiencies is minimized. One such concept, a hybrid system in which an ejector driven by engine reject heat is used to increase the performance of an internal combustion engine-driven heat pump, was analyzed by modifying an existing ejector heat pump model and combining it with generic compressor and internal combustion engine models. Under the model assumptions for nominal cooling mode conditions, the results showed that hybrid systems could providemore » substantial performance augmentation/emdash/up to 17/percent/ increase in system coefficient of performance for a parallel arrangement of an enhanced ejector with the engine-driven compressor. 4 refs., 4 figs., 4 tabs.« less

Calculation of Supersonic Combustion Using Implicit Schemes

NASA Technical Reports Server (NTRS)

Yoon, Seokkwan; Kwak, Dochan (Technical Monitor)

2003-01-01

One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle fleet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high specific impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors.

Combustion and Magnetohydrodynamic Processes in Advanced Pulse Detonation Rocket Engines

NASA Astrophysics Data System (ADS)

Cole, Lord Kahil

A number of promising alternative rocket propulsion concepts have been developed over the past two decades that take advantage of unsteady combustion waves in order to produce thrust. These concepts include the Pulse Detonation Rocket Engine (PDRE), in which repetitive ignition, propagation, and reflection of detonations and shocks can create a high pressure chamber from which gases may be exhausted in a controlled manner. The Pulse Detonation Rocket Induced Magnetohydrodynamic Ejector (PDRIME) is a modification of the basic PDRE concept, developed by Cambier (1998), which has the potential for performance improvements based on magnetohydrodynamic (MHD) thrust augmentation. The PDRIME has the advantage of both low combustion chamber seeding pressure, per the PDRE concept, and efficient energy distribution in the system, per the rocket-induced MHD ejector (RIME) concept of Cole, et al. (1995). In the initial part of this thesis, we explore flow and performance characteristics of different configurations of the PDRIME, assuming quasi-one-dimensional transient flow and global representations of the effects of MHD phenomena on the gas dynamics. By utilizing high-order accurate solvers, we thus are able to investigate the fundamental physical processes associated with the PDRIME and PDRE concepts and identify potentially promising operating regimes. In the second part of this investigation, the detailed coupling of detonations and electric and magnetic fields are explored. First, a one-dimensional spark-ignited detonation with complex reaction kinetics is fully evaluated and the mechanisms for the different instabilities are analyzed. It is found that complex kinetics in addition to sufficient spatial resolution are required to be able to quantify high frequency as well as low frequency detonation instability modes. Armed with this quantitative understanding, we then examine the interaction of a propagating detonation and the applied MHD, both in one-dimensional and two-dimensional transient simulations. The dynamics of the detonation are found to be affected by the application of magnetic and electric fields. We find that the regularity of one-dimensional cesium-seeded detonations can be significantly altered by reasonable applied magnetic fields (Bz ≤ 8T), but that it takes a stronger applied field (Bz > 16T) to significantly alter the cellular structure and detonation velocity of a two-dimensional detonation in the time in which these phenomena were observed. This observation is likely attributed to the additional coupling of the two-dimensional detonation with the transverse waves, which are not captured in the one-dimensional simulations. Future studies involving full ionization kinetics including collisional-radiative processes, will be used to examine these processes in further detail.

A 1/10 Scale Model Test of a Fixed Chute Mixer-Ejector Nozzle in Unsuppressed Model. Part 1; Test Overview

NASA Technical Reports Server (NTRS)

Wolter, John D.

2007-01-01

This paper discusses a test of a nozzle concept for a high-speed commercial aircraft. While a great deal of effort has been expended to und erstand the noise-suppressed, take-off performance of mixer-ejector n ozzles, little has been done to assess their performance in unsuppressed mode at other flight conditions. To address this, a 1/10th scale m odel mixer-ejector nozzle in unsuppressed mode was tested at conditio ns representing transonic acceleration, supersonic cruise, subsonic cruise, and approach. Various configurations were tested to understand the effects of acoustic liners and several geometric parameters, such as throat area, expansion ratio, and nozzle length on nozzle performance. Thrust, flow, and internal pressures were measured. A statistica l model of the peak thrust coefficient results is presented and discussed.

38th JANNAF Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Eggleston, Debra S. (Editor); Gannaway, Mary T. (Editor)

2002-01-01

This volume, the first of two volumes, is a collection of 55 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 38th Combustion Subcommittee (CS), 26 th Airbreathing Propulsion Subcommittee (APS), 20th Propulsion Systems Hazards Subcommittee (PSHS), and 21 Modeling and Simulation Subcommittee. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics cover five major technology areas including: 1) Combustion - Propellant Combustion, Ingredient Kinetics, Metal Combustion, Decomposition Processes and Material Characterization, Rocket Motor Combustion, and Liquid & Hybrid Combustion; 2) Liquid Rocket Engines - Low Cost Hydrocarbon Liquid Rocket Engines, Liquid Propulsion Turbines, Liquid Propulsion Pumps, and Staged Combustion Injector Technology; 3) Modeling & Simulation - Development of Multi- Disciplinary RBCC Modeling, Gun Modeling, and Computational Modeling for Liquid Propellant Combustion; 4) Guns Gun Propelling Charge Design, and ETC Gun Propulsion; and 5) Airbreathing - Scramjet an Ramjet- S&T Program Overviews.

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 1; Steady Predictions

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Graham, Jason S.; Ahuja, Vineet; Hosangadi, Ashvin

2010-01-01

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively.

Wall roughness effect on gas dynamics in supersonic ejector

NASA Astrophysics Data System (ADS)

Aronson, K. E.; Brezgin, D. V.

2016-10-01

The paper presents the numerical simulations results in order to figure out the influence of the wall surface roughness on gas-dynamic processes inside the supersonic ejector. For these purposes two commercial CFD-solvers (Star-CCM+ and Fluent) were used. A detailed comparative study of the built-in tools and approaches in both CFD-packages for evaluation of surface roughness effects on the logarithmic law velocity distribution inside the boundary layer is carried out. Influence of ejector surface roughness is compared with the influence of the backpressure. It is found out that either increasing the backpressure behind the ejector or increasing the surface roughness height, the appearance section of a pressure shock is displaced upstream (closer to the primary nozzle). The numerical simulations results of the ejector with rough walls in both CFD-solvers are well quantitative agreed between each other in terms of the mass flow rates and are well qualitative consistent in terms of the local flow parameters distribution. It is found out that in case of exceeding the "critical roughness height" for the given geometry and boundary conditions, the ejector switches to the "off-design" mode and its performance is significantly reduced.

The surface roughness effect on the performance of supersonic ejectors

NASA Astrophysics Data System (ADS)

Brezgin, D. V.; Aronson, K. E.; Mazzelli, F.; Milazzo, A.

2017-07-01

The paper presents the numerical simulation results of the surface roughness influence on gas-dynamic processes inside flow parts of a supersonic ejector. These simulations are performed using two commercial CFD solvers (Star- CCM+ and Fluent). The results are compared to each other and verified by a full-scale experiment in terms of global flow parameters (the entrainment ratio: the ratio between secondary to primary mass flow rate - ER hereafter) and local flow parameters distribution (the static pressure distribution along the mixing chamber and diffuser walls). A detailed comparative study of the employed methods and approaches in both CFD packages is carried out in order to estimate the roughness effect on the logarithmic law velocity distribution inside the boundary layer. Influence of the surface roughness is compared with the influence of the backpressure (static pressure at the ejector outlet). It has been found out that increasing either the ejector backpressure or the surface roughness height, the shock position displaces upstream. Moreover, the numerical simulation results of an ejector with rough walls in the both CFD solvers are well quantitatively agreed with each other in terms of the mean ER and well qualitatively agree in terms of the local flow parameters distribution. It is found out that in the case of exceeding the "critical roughness height" for the given boundary conditions and ejector's geometry, the ejector switches to the "off-design" mode and its performance decreases considerably.

Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

NASA Technical Reports Server (NTRS)

Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

2002-01-01

The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjet/scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

NASA Technical Reports Server (NTRS)

Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

2003-01-01

The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic air- breathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demon- strate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

Multidisciplinary Analysis of a Hypersonic Engine

NASA Technical Reports Server (NTRS)

Suresh, Ambady; Stewart, Mark

2003-01-01

The objective is to develop high fidelity tools that can influence ISTAR design In particular, tools for coupling Fluid-Thermal-Structural simulations RBCC/TBCC designers carefully balance aerodynamic, thermal, weight, & structural considerations; consistent multidisciplinary solutions reveal details (at modest cost) At Scram mode design point, simulations give details of inlet & combustor performance, thermal loads, structural deflections.

Development of an Experiment High Performance Nozzle Research Program

NASA Technical Reports Server (NTRS)

2004-01-01

As proposed in the above OAI/NASA Glenn Research Center (GRC) Co-Operative Agreement the objective of the work was to provide consultation and assistance to the NASA GRC GTX Rocket Based Combined Cycle (RBCC) Program Team in planning and developing requirements, scale model concepts, and plans for an experimental nozzle research program. The GTX was one of the launch vehicle concepts being studied as a possible future replacement for the aging NASA Space Shuttle, and was one RBCC element in the ongoing NASA Access to Space R&D Program (Reference 1). The ultimate program objective was the development of an appropriate experimental research program to evaluate and validate proposed nozzle concepts, and thereby result in the optimization of a high performance nozzle for the GTX launch vehicle. Included in this task were the identification of appropriate existing test facilities, development of requirements for new non-existent test rigs and fixtures, develop scale nozzle model concepts, and propose corresponding test plans. Also included were the evaluation of originally proposed and alternate nozzle designs (in-house and contractor), evaluation of Computational Fluid Dynamics (CFD) study results, and make recommendations for geometric changes to result in improved nozzle thrust coefficient performance (Cfg).

Experimental and Numerical Study on Supersonic Ejectors Working with R-1234ze(E)

NASA Astrophysics Data System (ADS)

Kracik, Jan; Dvorak, Vaclav; Nguyen Van, Vu; Smierciew, Kamil

2018-06-01

These days, much effort is being put into lowering the consumption of electric energy and involving renewable energy sources. Many engineers and designers are trying to develop environment-friendly technologies worldwide. It is related to incorporating appropriate devices into such technologies. The object of this paper is to investigate these devices in connection with refrigeration systems. Ejectors can be considered such as these devices. The primary interest of this paper is to investigate the suitability of a numerical model for an ejector, which is incorporated into a refrigeration system. In the present paper, there have been investigated seven different test runs of working of the ejector with a working fluid R-1234ze(E). Some of the investigated cases seem to have a good agreement and there are no significant discrepancies between them, however, there are also cases that do not correspond to the experimental data at all. The ejector has been investigated in both on-design and off-design working modes. A comparison between the experimental and numerical data (CFD) performed by Ansys Fluent software is presented and discussed for both an ideal and a real gas model. In addition, an enhanced analytical model has been introduced for all runs of the ejector.

A3 Subscale Rocket Hot Fire Testing

NASA Technical Reports Server (NTRS)

Saunders, G. P.; Yen, J.

2009-01-01

This paper gives a description of the methodology and results of J2-X Subscale Simulator (JSS) hot fire testing supporting the A3 Subscale Diffuser Test (SDT) project at the E3 test facility at Stennis Space Center, MS (SSC). The A3 subscale diffuser is a geometrically accurate scale model of the A3 altitude simulating rocket test facility. This paper focuses on the methods used to operate the facility and obtain the data to support the aerodynamic verification of the A3 rocket diffuser design and experimental data quantifying the heat flux throughout the facility. The JSS was operated at both 80% and 100% power levels and at gimbal angle from 0 to 7 degrees to verify the simulated altitude produced by the rocket-rocket diffuser combination. This was done with various secondary GN purge loads to quantify the pumping performance of the rocket diffuser. Also, special tests were conducted to obtain detailed heat flux measurements in the rocket diffuser at various gimbal angles and in the facility elbow where the flow turns from vertical to horizontal upstream of the 2nd stage steam ejector.

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 1; Steady Predictions

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Graham, Jason S.; Ahuja, Vineet; Hosangadi, Ashvin

2008-01-01

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

Spaceliner Class Operability Gains Via Combined Airbreathing/ Rocket Propulsion: Summarizing an Operational Assessment of Highly Reusable Space Transports

NASA Technical Reports Server (NTRS)

Nix, Michael B.; Escher, William J. d.

1999-01-01

In discussing a new NASA initiative in advanced space transportation systems and technologies, the Director of the NASA Marshall Space Flight Center, Arthur G. Stephenson, noted that, "It would use new propulsion technology, air-breathing engine so you don't have to carry liquid oxygen, at least while your flying through the atmosphere. We are calling it Spaceliner 100 because it would be 100 times cheaper, costing $ 100 dollars a pound to orbit." While airbreathing propulsion is directly named, rocket propulsion is also implied by, "... while you are flying through the atmosphere." In-space final acceleration to orbital speed mandates rocket capabilities. Thus, in this informed view, Spaceliner 100 will be predicated on combined airbreathing/rocket propulsion, the technical subject of this paper. Interestingly, NASA's recently concluded Highly Reusable Space Transportation (HRST) study focused on the same affordability goal as that of the Spaceliner 100 initiative and reflected the decisive contribution of combined propulsion as a way of expanding operability and increasing the design robustness of future space transports, toward "aircraft like" capabilities. The HRST study built on the Access to Space Study and the Reusable Launch Vehicle (RLV) development activities to identify and characterize space transportation concepts, infrastructure and technologies that have the greatest potential for reducing delivery cost by another order of magnitude, from $1,000 to $100-$200 per pound for 20,000 lb. - 40.000 lb. payloads to low earth orbit (LEO). The HRST study investigated a number of near-term, far-term, and very far-term launch vehicle concepts including all-rocket single-stage-to-orbit (SSTO) concepts, two-stage-to-orbit (TSTO) concepts, concepts with launch assist, rocket-based combined cycle (RBCC) concepts, advanced expendable vehicles, and more far term ground-based laser powered launchers. The HRST study consisted of preliminary concept studies, assessments and analysis tool development for advanced space transportation systems, followed by end-to-end system concept definitions and trade analyses, specific system concept definition and analysis, specific key technology and topic analysis, system, operational and economics model development, analysis, and integrated assessments. The HRST Integration Task Force (HITF) was formed to synthesize study results in several specific topic areas and support the development of conclusions from the study: Systems Concepts Definitions, Technology Assessment, Operations Assessment, and Cost Assessment. This paper summarizes the work of the Operations Assessment Team: the six approaches used, the analytical tools and methodologies developed and employed, the issues and concerns, and the results of the assessment. The approaches were deliberately varied in measures of merit and procedure to compensate for the uncertainty inherent in operations data in this early phase of concept exploration. In general, rocket based combined cycle (RBCC) concepts appear to have significantly greater potential than all-rocket concepts for reducing operations costs.

Self-focused acoustic ejectors for viscous liquids.

PubMed

Hon, S F; Kwok, K W; Li, H L; Ng, H Y

2010-06-01

Self-focused acoustic ejectors using the Fresnel zone plate (FZP) have been developed for ejecting viscous liquids, without nozzle, in the drop-on-demand mode. The FZP is composed of a lead zirconate titanate piezoelectric plate patterned with a series of annular electrodes, with the unelectroded region of the plate removed. Our results show that the acoustic waves are effectively self-focused by constructive interference in glycerin (with a viscosity of 1400 mPa s), giving small focal points with a high pressure. Due to the high attenuation, the wave pressure decreases significantly with the distance from the FZP. Nevertheless, the pressure at the focal points 2.5 and 6.5 mm from the FZP is high enough to eject glycerin droplets in the drop-on-demand mode. Driven by a simple wave train comprising a series of sinusoidal voltages with an amplitude of 35 V, a frequency of 4.28 MHz, and a duration of 2 ms, the ejector can eject fine glycerin droplets with a diameter of 0.4 mm at a repetition frequency of 120 Hz in a downward direction. Droplets of other viscous liquids, such as the prepolymer of an epoxy with a viscosity of 2000 mPa s, can also be ejected in the drop-on-demand mode under similar conditions.

Studies of Operating Frequency Effects On Ejector-based Thrust Augmentation in a Pulse Detonation Engine

NASA Technical Reports Server (NTRS)

Landry, K.

2005-01-01

Studies were performed in order to characterize the thrust augmentation potential of an ejector in a Pulse Detonation Engine application. A 49-mm diameter tube of 0.914-m length was constructed with one open end and one closed end. Ethylene, oxygen, and nitrogen were introduced into the tube at the closed end through the implementation of a fast mixing injector. The tube was completely filled with a stoichiometric mixture containing a one to one molar ratio of nitrogen to oxygen. Ethylene was selected as the fuel due to its detonation sensitivity and the molar ratio of the oxidizer was chosen for heat transfer purposes. Detonations were initiated in the tube through the use of a spark ignition system. The PDE was operated in a multi-cycle mode at frequencies ranging from 20-Hz to 50-Hz. Baseline thrust measurements with no ejector present were performed while operating the engine at various frequencies and compared to theoretical estimates. The baseline values were observed to agree with the theoretical model at low operating frequencies and proved to be increasingly lower than the predicted values as the operating frequency was increased. The baseline thrust measurements were observed to agree within 15 percent of the model for all operating frequencies. A straight 152-mm diameter ejector was installed and thrust augmentation percentages were measured. The length of the ejector was varied while the overlap percentage (percent of the ejector length which overlapped the tube) was maintained at 25 percent for all tests. In addition, the effect of ejector inlet geometry was investigated by comparing results with a straight inlet to those of a 38-mm inlet diameter. The thrust augmentation of the straight inlet ejector proved to be independent of engine operating frequency, augmenting thrust by 40 percent for the 0.914-m length ejector. In contrast, the rounded lip ejector of the same length seemed to be highly dependent on the engine operating frequency. An optimum operating frequency observed with the rounded inlet occurred at an operating frequency of 30-Hz, resulting in thrust augmentation percentages greater than 100 percent. The effect that the engine operating frequency had on thrust augmentation levels attained with an ejector was characterized and optimum performance parameters were established. Insight into the frequency dependent nature of the ejector performance was pursued. Suggestions for future experiments which are needed to fully understand the means in which thrust augmentation is achieved in a PDE-ejector configuration were noted.

Preliminary Sizing Completed for Single- Stage-To-Orbit Launch Vehicles Powered By Rocket-Based Combined Cycle Technology

NASA Technical Reports Server (NTRS)

Roche, Joseph M.

2002-01-01

Single-stage-to-orbit (SSTO) propulsion remains an elusive goal for launch vehicles. The physics of the problem is leading developers to a search for higher propulsion performance than is available with all-rocket power. Rocket-based combined cycle (RBCC) technology provides additional propulsion performance that may enable SSTO flight. Structural efficiency is also a major driving force in enabling SSTO flight. Increases in performance with RBCC propulsion are offset with the added size of the propulsion system. Geometrical considerations must be exploited to minimize the weight. Integration of the propulsion system with the vehicle must be carefully planned such that aeroperformance is not degraded and the air-breathing performance is enhanced. Consequently, the vehicle's structural architecture becomes one with the propulsion system architecture. Geometrical considerations applied to the integrated vehicle lead to low drag and high structural and volumetric efficiency. Sizing of the SSTO launch vehicle (GTX) is itself an elusive task. The weight of the vehicle depends strongly on the propellant required to meet the mission requirements. Changes in propellant requirements result in changes in the size of the vehicle, which in turn, affect the weight of the vehicle and change the propellant requirements. An iterative approach is necessary to size the vehicle to meet the flight requirements. GTX Sizer was developed to do exactly this. The governing geometry was built into a spreadsheet model along with scaling relationships. The scaling laws attempt to maintain structural integrity as the vehicle size is changed. Key aerodynamic relationships are maintained as the vehicle size is changed. The closed weight and center of gravity are displayed graphically on a plot of the synthesized vehicle. In addition, comprehensive tabular data of the subsystem weights and centers of gravity are generated. The model has been verified for accuracy with finite element analysis. The final trajectory was rerun using OTIS (Boeing Corporation's trajectory optimization software package), and the sizing output was incorporated into a solid model of the vehicle using PRO/Engineer computer-aided design software (Parametric Technology Corporation, Waltham, MA).

Magnetohydrodynamic Augmentation of Pulse Detonation Rocket Engines (Preprint)

DTIC Science & Technology

2010-09-28

augmentation of the thrust . Ejectors typically transfer energy between streams through shear stress between separate flow streams, where a portion of the...the opportunity to extract energy and apply it to a separate stream where the net thrust can be increased. With MHD augmentation , such as in the Pulse...with the PDRIME for separate or additional thrust augmentation . Results show potential performance gains under many flight and operating conditions

One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

NASA Technical Reports Server (NTRS)

Sakowski, Barbara; Edwards, Daryl; Dickens, Kevin

2014-01-01

Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation [Ref 1]. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet conduction as well as the degrading effect of mass and heat transfer due to the presence of noncondensibles. The one dimension model of the condensing spray chamber makes no presupposition on the pressure profile within the chamber, allowing the implemented droplet physics of heat and mass transfer coupled to the SINDAFLUINT solver to determine a transient pressure profile of the condensing spray chamber. Model results compare well to the RL-10 engine pressure test data.

Hetero-oligomerization among the TIF family of RBCC/TRIM domain-containing nuclear cofactors: a potential mechanism for regulating the switch between coactivation and corepression.

PubMed

Peng, Hongzhuang; Feldman, Irina; Rauscher, Frank J

2002-07-12

The RING-B box-coiled-coil (RBCC) motif (also re-named recently as the tripartite motif (TRIM)) is a widely distributed motif that is hypothesized to be a protein-protein interface. The RBCC/TRIM domain of the corepressor KAP-1 is both necessary and sufficient to interact directly with the transcription repressor KRAB domain. Each subdomain of the KAP-1-RBCC contributes directly to the oligomerization and/or ligand recognition. Little is known about the function or the natural binding ligands for the RBCC/TRIM domain of the other TIF family members. In order to investigate whether hetero-oligomerization might be a biological regulatory mechanism, we have evaluated the hetero-oligomerization potential of the TIF family members including KAP-1, TIF1alpha, TIF1gamma, and the RBCC/TRIM family members including PML1, and MID1. We have reconstituted and characterized the oligomerization for these proteins using baculovirus and mammalian expression systems by biochemical approaches. Our data indicate that the RBCC/TRIM domains of KAP-1, TIF1alpha and TIF1gamma exist in a homo-oligomeric state. However, there is little cross-talk between KAP-1 and other TIF family members, suggesting that a high degree of specificity for oligomerization interface and ligand recognition is intrinsically built into the RBCC/TRIM domain of KAP-1. Finally, we demonstrate that TIF1alpha interacts with TIF1gamma and the coiled-coil region of TIF1gamma is necessary for this interaction. The hetero-oligomerization between TIF1alpha and TIF1gamma implies a potential regulatory mechanism for transcriptional regulation. (c) 2002 Elsevier Science Ltd.

Marquardt's Mach 4.5 Supercharged Ejector Ramjet (SERJ) High-Performance Aircraft Engine Project: Unfulfilled Aspirations Ca.1970

NASA Technical Reports Server (NTRS)

Escher, William J. D.; Roddy, Jordan E.; Hyde, Eric H.

2000-01-01

The Supercharged Ejector Ramjet (SERJ) engine developments of the 1960s, as pursued by The Marquardt Corporation and its associated industry team members, are described. In just three years, engineering work on this combined-cycle powerplant type evolved, from its initial NASA-sponsored reusable space transportation system study status, into a U.S. Air Force/Navy-supported exploratory development program as a candidate 4.5 high-performance military aircraft engine. Bridging a productive transition from the spaceflight to the aviation arena, this case history supports the expectation that fully-integrated airbreathing/rocket propulsion systems hold high promise toward meeting the demanding propulsion requirements of tomorrow's aircraft-like Spaceliner class transportation systems. Lessons to be learned from this "SERJ Story" are offered for consideration by today's advanced space transportation and combined-cycle propulsion researchers and forward-planning communities.

Numerical Simulation of Pulse Detonation Rocket-Induced MHD Ejector (PDRIME) Concepts for Advanced Propulsion Systems

DTIC Science & Technology

2012-02-28

Coupling in Detonation Waves: 1D Dynamics”, Paper 89, 23rd International Colloquium on the Dynamics of Explosions and Reactive ...and temperature, and can be modeled as a constant volume reaction , which is more efficient than a constant pressure reaction . After the detonation ... kinetics , and flow processes using high order numerical methods. A fifth-order WENO (weighted essentially non -oscillatory12,13) scheme was used

Launch Vehicle Systems Analysis

NASA Technical Reports Server (NTRS)

Olds, John R.

1999-01-01

This report summaries the key accomplishments of Georgia Tech's Space Systems Design Laboratory (SSDL) under NASA Grant NAG8-1302 from NASA - Marshall Space Flight Center. The report consists of this summary white paper, copies of technical papers written under this grant, and several viewgraph-style presentations. During the course of this grant four main tasks were completed: (1)Simulated Combined-Cycle Rocket Engine Analysis Module (SCCREAM), a computer analysis tool for predicting the performance of various RBCC engine configurations; (2) Hyperion, a single stage to orbit vehicle capable of delivering 25,000 pound payloads to the International Space Station Orbit; (3) Bantam-X Support - a small payload mission; (4) International Trajectory Support for interplanetary human Mars missions.

Improvement to the Standardization Life-Cycle-Cost Model for Hull, Mechanical, and Electrical Components and Equipments.

DTIC Science & Technology

1980-03-01

APL 061900283, Compressor, Motor , AC, 440 V, Valve, Solenoid Centrifugal, 230 gpm Air 300 hp Each C/E was identified to a distinguishable complexity...Deck and Hull Machinery D Rocket Handling Agitator-Paint Shaker Deck and Hull Machinery B Air Conditioner Refrigeration/Heating C Systems Air Ejector...Terminal Electrical Systems a Brake- Air Deck and Hull Machinery D Brake-Electric, Motor Operated Deck and Hull Machinery D Srake-Electric, Solenoid

Advanced Technology Inlet Design, NRA 8-21 Cycle II: DRACO Flowpath Hypersonic Inlet Design

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

1999-01-01

The report outlines work performed in support of the flowpath development for the DRACO engine program. The design process initiated to develop a hypersonic axisymmetric inlet for a Mach 6 rocket-based combined cycle (RBCC) engine is discussed. Various design parametrics were investigated, including design shock-on-lip Mach number, cone angle, throat Mach number, throat angle. length of distributed compression, and subsonic diffuser contours. Conceptual mechanical designs consistent with installation into the D-21 vehicle were developed. Additionally, program planning for an intensive inlet development program to support a Critical Design Review in three years was performed. This development program included both analytical and experimental elements and support for a flight-capable inlet mechanical design.

Credit BG. View looking northeast at southwestern side of Test ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. View looking northeast at southwestern side of Test Stand "D" complex. Test Stand "D" workshop (Building 4222/E-23) is at left; shed to its immediate right is an entrance to underground tunnel system which interconnects all test stands. To the right of Test Stand "D" tower are four Clayton water-tube flash boilers once used in the Steam Generator Plant 4280/E-81 to power the vacuum ejector system at "D" and "C" stands. A corner of 4280/E-81 appears behind the boilers. Boilers were removed as part of stand dismantling program. The Dv (vertical vacuum) Test Cell is located in the Test Stand "D" tower, behind the sunscreen on the west side. The top of the tower contains a hoist for lifting or lowering rocket engines into the Dv Cell. Other equipment mounted in the tower is part of the steam-driven vacuum ejector system - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Credit BG. Looking southeast at Test Stand "D" (Building 4223/E24). ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. Looking southeast at Test Stand "D" (Building 4223/E-24). Left foreground contains six high-pressure nitrogen tanks which supplied nitrogen for operation of propellant valves. Several tanks for other substances have been removed from the base of the tower as part of decontamination and dismantling program. The vertical vacuum test cell can be seen in the tower behind the western sunscreen. At the top of the tower in the northeast corner is the interstage condenser used in the series of vacuum ejectors; at the top of the condenser is one of two Z-stage ejectors used to evacuate the condenser. The hoist beam for lifting/lowering rocket engines can be clearly seen projecting to the west over the pavement. In the distance on the right are Clayton water-tube steam generators from Building 4280/E-81, and the towers for Test Stand "C" and its scrubber-condenser - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

An affordable RBCC-powered 2-stage small orbital payload transportation systems concept based on test-proven hardware

NASA Astrophysics Data System (ADS)

Escher, William J. D.

1998-01-01

Deriving from the initial planning activity of early 1965, which led to NASA's Advanced Space Transportation Program (ASTP), an early-available airbreathing/rocket combined propulsion system powered ``ultralight payload'' launcher was defined at the conceptual design level. This system, named the ``W Vehicle,'' was targeted to be a ``second generation'' successor to the original Bantam Lifter class, all-rocket powered systems presently being pursued by NASA and a selected set of its contractors. While this all-rocket vehicle is predicated on a fully expendable approach, the W-Vehicle system was to be a fully reusable 2-stage vehicle. The general (original) goal of the Bantam class of launchers was to orbit a 100 kg payload for a recurring per-launch cost of less than one million dollars. Reusability, as the case for larger vehicles focusing on single stage to orbit (SSTO) configurations, is considered the principal key to affordability. In the general context of a range of space transports, covering the payload range of 0.1 to 10 metric ton payloads, the W Vehicle concept-predicated mainly on ground- and flight-test proven hardware-is described in this paper, along with a nominal development schedule and budgetary estimate (recurring costs were not estimated).

B-1 and B-3 Test Stands at NASA’s Plum Brook Station

NASA Image and Video Library

1966-09-21

Operation of the High Energy Rocket Engine Research Facility (B-1), left, and Nuclear Rocket Dynamics and Control Facility (B-3) at the National Aeronautics and Space Administration’s (NASA) Plum Brook Station in Sandusky, Ohio. The test stands were constructed in the early 1960s to test full-scale liquid hydrogen fuel systems in simulated altitude conditions. Over the next decade each stand was used for two major series of liquid hydrogen rocket tests: the Nuclear Engine for Rocket Vehicle Application (NERVA) and the Centaur second-stage rocket program. The different components of these rocket engines could be studied under flight conditions and adjusted without having to fire the engine. Once the preliminary studies were complete, the entire engine could be fired in larger facilities. The test stands were vertical towers with cryogenic fuel and steam ejector systems. B-1 was 135 feet tall, and B-3 was 210 feet tall. Each test stand had several levels, a test section, and ground floor shop areas. The test stands relied on an array of support buildings to conduct their tests, including a control building, steam exhaust system, and fuel storage and pumping facilities. A large steam-powered altitude exhaust system reduced the pressure at the exhaust nozzle exit of each test stand. This allowed B-1 and B-3 to test turbopump performance in conditions that matched the altitudes of space.

A Concept of Two-Stage-To-Orbit Reusable Launch Vehicle

NASA Astrophysics Data System (ADS)

Yang, Yong; Wang, Xiaojun; Tang, Yihua

2002-01-01

Reusable Launch Vehicle (RLV) has a capability of delivering a wide rang of payload to earth orbit with greater reliability, lower cost, more flexibility and operability than any of today's launch vehicles. It is the goal of future space transportation systems. Past experience on single stage to orbit (SSTO) RLVs, such as NASA's NASP project, which aims at developing an rocket-based combined-cycle (RBCC) airplane and X-33, which aims at developing a rocket RLV, indicates that SSTO RLV can not be realized in the next few years based on the state-of-the-art technologies. This paper presents a concept of all rocket two-stage-to-orbit (TSTO) reusable launch vehicle. The TSTO RLV comprises an orbiter and a booster stage. The orbiter is mounted on the top of the booster stage. The TSTO RLV takes off vertically. At the altitude about 50km the booster stage is separated from the orbiter, returns and lands by parachutes and airbags, or lands horizontally by means of its own propulsion system. The orbiter continues its ascent flight and delivers the payload into LEO orbit. After completing orbit mission, the orbiter will reenter into the atmosphere, automatically fly to the ground base and finally horizontally land on the runway. TSTO RLV has less technology difficulties and risk than SSTO, and maybe the practical approach to the RLV in the near future.

Performance tests of a two phase ejector

DOE Office of Scientific and Technical Information (OSTI.GOV)

Harrell, G.S.; Kornhauser, A.A.

1995-12-31

The ejector expansion refrigeration cycle is a modified vapor compression cycle in which a two phase ejector is used to recover a portion of the work otherwise lost in the expansion valve. The ejector improves cycle performance by increasing compressor inlet pressure and by lowering the quality of liquid entering the evaporator. Theoretically, a cooling COP improvement of approximately 23% is achievable for a typical refrigerating cycle and an ideal ejector. If the ejector performed as well as typical single phase ejectors an improvement of 12% could be achieved. Previous tests have demonstrated a smaller 3.7% improvement; the difference ismore » in the poor performance of the two phase ejector. The purpose of this research is to understand the operating characteristics of the two phase ejector and to devise design improvements. A two phase ejector test rig has been constructed and tested. Preliminary data show performance superior to previously tested two phase ejectors, but still inferior to single phase ejectors. Ejector performance corresponds to refrigeration cycle COP improvements ranging from 3.9% to 7.6%.« less

Review on factors affecting the performance of pulse detonation engine

NASA Astrophysics Data System (ADS)

Tripathi, Saurabh; Pandey, Krishna Murari

2018-04-01

Now a day's rocket engines (air-breathing type) are being used for aerospace purposes but the studies have shown that these are less efficient, so alternatives are being searched for these. Pulse Detonation Engine (PDE) is one such efficient engine which can replace the rocket engines. In this review paper, different researches have been cited. As can be observed from various researches, insertion of obstacles is better. Deflagration to Detonation(DDT) transition process is found to be most important factor. So a lot of researches are being done considering this DDT chamber. Also, the ignition chamber and ejector were found to improve the effectiveness of PDE. The PDE works with a range of Mach 0-4. Flame acceleration is also found to increase the DDT process. Use of valve and valveless engine has also been compared. Various other factors have been focused in this review paper which is found to boost PDE performance.

Exchange inlet optimization by genetic algorithm for improved RBCC performance

NASA Astrophysics Data System (ADS)

Chorkawy, G.; Etele, J.

2017-09-01

A genetic algorithm based on real parameter representation using a variable selection pressure and variable probability of mutation is used to optimize an annular air breathing rocket inlet called the Exchange Inlet. A rapid and accurate design method which provides estimates for air breathing, mixing, and isentropic flow performance is used as the engine of the optimization routine. Comparison to detailed numerical simulations show that the design method yields desired exit Mach numbers to within approximately 1% over 75% of the annular exit area and predicts entrained air massflows to between 1% and 9% of numerically simulated values depending on the flight condition. Optimum designs are shown to be obtained within approximately 8000 fitness function evaluations in a search space on the order of 106. The method is also shown to be able to identify beneficial values for particular alleles when they exist while showing the ability to handle cases where physical and aphysical designs co-exist at particular values of a subset of alleles within a gene. For an air breathing engine based on a hydrogen fuelled rocket an exchange inlet is designed which yields a predicted air entrainment ratio within 95% of the theoretical maximum.

Propulsion Research and Technology: Overview

NASA Technical Reports Server (NTRS)

Cole, John; Schmidt, George

1999-01-01

Propulsion is unique in being the main delimiter on how far and how fast one can travel in space. It is the lack of truly economical high-performance propulsion systems that continues to limit and restrict the extent of human endeavors in space. Therefore the goal of propulsion research is to conceive and investigate new, revolutionary propulsion concepts. This presentation reviews the development of new propulsion concepts. Some of these concepts are: (1) Rocket-based Combined Cycle (RBCC) propulsion, (2) Alternative combined Cycle engines suc2 as the methanol ramjet , and the liquid air cycle engines, (3) Laser propulsion, (4) Maglifter, (5) pulse detonation engines, (6) solar thermal propulsion, (7) multipurpose hydrogen test bed (MHTB) and other low-G cryogenic fluids, (8) Electric propulsion, (9) nuclear propulsion, (10) Fusion Propulsion, and (11) Antimatter technology. The efforts of the NASA centers in this research is also spotlighted.

Rocket Engines Displayed for 1966 Inspection at Lewis Research Center

NASA Image and Video Library

1966-10-21

An array of rocket engines displayed in the Propulsion Systems Laboratory for the 1966 Inspection held at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis engineers had been working on chemical, nuclear, and solid rocket engines throughout the 1960s. The engines on display are from left to right: two scale models of the Aerojet M-1, a Rocketdyne J-2, a Pratt and Whitney RL-10, and a Rocketdyne throttleable engine. Also on display are several ejector plates and nozzles. The Chemical Rocket Division resolved issues such as combustion instability and screech, and improved operation of cooling systems and turbopumps. The 1.5-million pound thrust M-1 engine was the largest hydrogen-fueled rocket engine ever created. It was a joint project between NASA Lewis and Aerojet-General. Although much larger in size, the M-1 used technology developed for the RL-10 and J-2. The M-1 program was cancelled in late 1965 due to budget cuts and the lack of a post-Apollo mission. The October 1966 Inspection was the culmination of almost a year of events held to mark the centers’ 25th anniversary. The three‐day Inspection, Lewis’ first since 1957, drew 2000 business, industry, and government executives and included an employee open house. The visitors witnessed presentations at the major facilities and viewed the Gemini VII spacecraft, a Centaur rocket, and other displays in the hangar. In addition, Lewis’ newest facility, the Zero Gravity Facility, was shown off for the first time.

The Abstraction Process of Limit Knowledge

ERIC Educational Resources Information Center

Sezgin Memnun, Dilek; Aydin, Bünyamin; Özbilen, Ömer; Erdogan, Günes

2017-01-01

The RBC+C abstraction model is an effective model in mathematics education because it gives the opportunity to analyze research data through cognitive actions. For this reason, we aim to examine the abstraction process of the limit knowledge of two volunteer participant students using the RBC+C abstraction model. With this aim, the students'…

D-21B RBCC Modification Feasibility Study

NASA Technical Reports Server (NTRS)

1999-01-01

This report presents a feasibility study on the modifications required to re-engine the Lockheed D-21 Drone for use as a NASA RBCC engine. An introduction, background information, engine configuration and performance, propulsion system integration, loads/thermal analysis, avionics/systems, flight test results, costs and work schedule, and some conclusions are presented.

The effect of proteins from animal source foods on heme iron bioavailability in humans.

PubMed

Pizarro, Fernando; Olivares, Manuel; Valenzuela, Carolina; Brito, Alex; Weinborn, Valerie; Flores, Sebastián; Arredondo, Miguel

2016-04-01

Forty-five women (35-45 year) were randomly assigned to three iron (Fe) absorption sub-studies, which measured the effects of dietary animal proteins on the absorption of heme Fe. Study 1 was focused on heme, red blood cell concentrate (RBCC), hemoglobin (Hb), RBCC+beef meat; study 2 on heme, heme+fish, chicken, and beef; and study 3 on heme and heme+purified animal protein (casein, collagen, albumin). Study 1: the bioavailability of heme Fe from Hb was similar to heme only (∼13.0%). RBCC (25.0%) and RBCC+beef (21.3%) were found to be increased 2- and 1.6-fold, respectively, when compared with heme alone (p<0.05). Study 2: the bioavailability from heme alone (10.3%) was reduced (p<0.05) when it was blended with fish (7.1%) and chicken (4.9%), however it was unaffected by beef. Study 3: casein, collagen, and albumin did not affect the bioavailability of Fe. Proteins from animal source foods and their digestion products did not enhance heme Fe absorption. Copyright © 2015. Published by Elsevier Ltd.

Parametric Investigation of Thrust Augmentation by Ejectors on a Pulsed Detonation Tube

NASA Technical Reports Server (NTRS)

Wilson, Jack; Sgondea, Alexandru; Paxson, Daniel E.; Rosenthal, Bruce N.

2006-01-01

A parametric investigation has been made of thrust augmentation of a 1 in. diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentation ratios for each ejector were fitted using a polynomial response surface, from which the optimum ratios of ejector diameter to detonation tube diameter, and ejector length and nose radius to ejector diameter, were found. Thrust augmentation ratios above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector.

Unsteady Ejector Performance: An Experimental Investigation Using a Resonance Tube Driver

NASA Technical Reports Server (NTRS)

Wilson, Jack; Paxson, Daniel E.

2002-01-01

A statistically designed experiment to characterize thrust augmentation for unsteady ejectors has been conducted at the NASA Glenn Research Center. The variable parameters included ejector diameter, length, and nose radius. The pulsed jet driving the ejectors was produced by a shrouded resonance (or Hartmann-Sprenger) tube. In contrast to steady ejectors, an optimum ejector diameter was found, which coincided with the diameter of the vortex ring created at the pulsed jet exit. Measurements of ejector exit velocity using a hot-wire permitted evaluation of the mass augmentation ratio, which was found to correlate to thrust augmentation following a formula derived for steady ejectors.

A new methodology for sizing and performance predictions of a rotary wing ejector

NASA Astrophysics Data System (ADS)

Moodie, Alex Montfort

The application of an ejector nozzle integrated with a reaction drive rotor configuration for a vertical takeoff and landing rotorcraft is considered in this research. The ejector nozzle is a device that imparts energy from a high speed airflow source to a lower speed secondary airflow inside a duct. The overall nozzle exhaust mass flow rate is increased through fluid entrainment, while the exhaust gas velocity is simultaneously decreased. The exhaust gas velocity is strongly correlated to the jet noise produced by the nozzle, making the ejector a good candidate for propulsion system noise reduction. Ejector nozzles are mechanically simple in that there are no moving parts. However, coupled fluid dynamic processes are involved, complicating analysis and design. Geometric definitions of the ejector nozzle are determined through a reduced fidelity, multi-disciplinary, representation of the rotary wing ejector. The resulting rotary wing ejector geometric sizing procedure relates standard vehicle and rotor design parameters to the ejector. Additionally, a rotary wing ejector performance procedure is developed to compare this rotor configuration to a conventional rotor. Performance characteristics and aerodynamic effects of the rotor and ejector nozzle are analytically studied. Ejector nozzle performance, in terms of exit velocities, is compared to the primary reaction drive nozzle; giving an indication of the potential for noise reduction. Computational fluid dynamics are paramount in predicting the aerodynamic effects of the ejector nozzle located at the rotor blade tip. Two-dimensional, steady-state, Reynolds-averaged Navier-Stokes (RANS) models are implemented for sectional lift and drag predictions required for the rotor aerodynamic model associated with both the rotary wing ejector sizing and performance procedures. A three-dimensional, unsteady, RANS simulation of the rotary wing ejector is performed to study the aerodynamic interactions between the ejector nozzle and rotor. Overall performance comparisons are made between the two- and three-dimensional models of the rotary wing ejector, and a similar conventional rotor.

Application of thrusting ejectors to tactical aircraft having vertical lift and short-field capability

NASA Technical Reports Server (NTRS)

Koenig, D. G.; Stoll, F.; Aoyagi, K.

1981-01-01

The status of ejector development in terms of application to V/STOL aircraft is reported in three categories: aircraft systems and ejector concepts; ejector performance including prediction techniques and experimental data base available; and, integration of the ejector with complete aircraft configurations. Available prediction techniques are reviewed and performance of three ejector designs with vertical lift capability is summarized. Applications of the 'fuselage' and 'short diffuser' ejectors to fighter aircraft are related to current and planned research programs. Recommendations are listed for effort needed to evaluate installed performance.

Development of a Liner Design Methodology and Relevant Results of Acoustic Suppression in the Farfield for Mixer-Ejector Nozzles

NASA Technical Reports Server (NTRS)

Salikuddin, M.

2006-01-01

We have developed a process to predict noise field interior to the ejector and in the farfield for any liner design for a mixer-ejector of arbitrary scale factor. However, a number of assumptions, not verified for the current application, utilized in this process, introduce uncertainties in the final result, especially, on a quantitative basis. The normal impedance model for bulk with perforated facesheet is based on homogeneous foam materials of low resistivity. The impact of flow conditions for HSCT application as well as the impact of perforated facesheet on predicted impedance is not properly accounted. Based on the measured normal impedance for deeper bulk samples (i.e., 2.0 in.) the predicted reactance is much higher compared to the data at frequencies above 2 kHz for T-foam and 200 ppi SiC. The resistance is under predicted at lower frequencies (below 4 kHz) for these samples. Thus, the use of such predicted data in acoustic suppression is likely to introduce inaccuracies. It should be noted that the impedance prediction methods developed recently under liner technology program are not utilized in the studies described in this report due to the program closeout. Acoustic suppression prediction is based on the uniform flow and temperature conditions in a two-sided treated constant area rectangular duct. In addition, assumptions of equal energy per mode noise field and interaction of all frequencies with the treated surface for the entire ejector length may not be accurate. While, the use of acoustic transfer factor minimizes the inaccuracies associated with the prediction for a known test case, the assumption of the same factor for other liner designs and with different linear scale factor ejectors seems to be very optimistic. As illustrated in appendix D that the predicted noise suppression for LSM-1 is lower compared to the measured data is an indication of the above argument. However, the process seems to be more reliable when used for the same scale models for different liner designs as demonstrated for Gen. 1 mixer-ejectors.

13. Coal ejectors mounted on aft bulkhead of coal bunker. ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

13. Coal ejectors mounted on aft bulkhead of coal bunker. Ejectors were used to flush overboard live coals and clinkers from firebed (pipe for carrying coals overboard has been removed from ejector in foreground). Coal doors from bunker appear beside ejector in foreground). Coal doors from bunker appear beside ejectors at deck; note firing shovels in background against hull. - Steamboat TICONDEROGA, Shelburne Museum Route 7, Shelburne, Chittenden County, VT

Performance Assessment of a Large Scale Pulsejet- Driven Ejector System

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Litke, Paul J.; Schauer, Frederick R.; Bradley, Royce P.; Hoke, John L.

2006-01-01

Unsteady thrust augmentation was measured on a large scale driver/ejector system. A 72 in. long, 6.5 in. diameter, 100 lb(sub f) pulsejet was tested with a series of straight, cylindrical ejectors of varying length, and diameter. A tapered ejector configuration of varying length was also tested. The objectives of the testing were to determine the dimensions of the ejectors which maximize thrust augmentation, and to compare the dimensions and augmentation levels so obtained with those of other, similarly maximized, but smaller scale systems on which much of the recent unsteady ejector thrust augmentation studies have been performed. An augmentation level of 1.71 was achieved with the cylindrical ejector configuration and 1.81 with the tapered ejector configuration. These levels are consistent with, but slightly lower than the highest levels achieved with the smaller systems. The ejector diameter yielding maximum augmentation was 2.46 times the diameter of the pulsejet. This ratio closely matches those of the small scale experiments. For the straight ejector, the length yielding maximum augmentation was 10 times the diameter of the pulsejet. This was also nearly the same as the small scale experiments. Testing procedures are described, as are the parametric variations in ejector geometry. Results are discussed in terms of their implications for general scaling of pulsed thrust ejector systems

Parallel Work of CO2 Ejectors Installed in a Multi-Ejector Module of Refrigeration System

NASA Astrophysics Data System (ADS)

Bodys, Jakub; Palacz, Michal; Haida, Michal; Smolka, Jacek; Nowak, Andrzej J.; Banasiak, Krzysztof; Hafner, Armin

2016-09-01

A performance analysis on of fixed ejectors installed in a multi-ejector module in a CO2 refrigeration system is presented in this study. The serial and the parallel work of four fixed-geometry units that compose the multi-ejector pack was carried out. The executed numerical simulations were performed with the use of validated Homogeneous Equilibrium Model (HEM). The computational tool ejectorPL for typical transcritical parameters at the motive nozzle were used in all the tests. A wide range of the operating conditions for supermarket applications in three different European climate zones were taken into consideration. The obtained results present the high and stable performance of all the ejectors in the multi-ejector pack.

Thrust Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pak, Sibtosh; Shehadeh, R.; Saretto, S. R.; Lee, S.-Y.

2005-01-01

Results of an experimental effort on pulse detonation driven ejectors aimed at probing different aspects of PDE ejector processes, are presented and discussed. The PDE was operated using ethylene as the fuel and an equimolar oxygen/nitrogen mixture as the oxidizer at an equivalence ratio of one. The thrust measurements for the PDE alone are in excellent agreement with experimental and modeling results reported in the literature and serve as a Baseline for the ejector studies. These thrust measurements were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups using constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Aeroacoustic Characteristics of a Rectangular Multi-Element Supersonic Jet Mixer-Ejector Nozzle

NASA Technical Reports Server (NTRS)

Raman, Ganesh; Taghavi, Ray

1996-01-01

This paper provides a unique, detailed evaluation of the acoustics and aerodynamics of a rectangular multi-element supersonic jet mixer-ejector noise suppressor. The performance of such mixer-ejectors is important in aircraft engine application for noise suppression and thrust augmentation. In contrast to most prior experimental studies on ejectors that reported either aerodynamic or acoustic data, our work documents both types of data. We present information on the mixing, pumping, ejector wall pressure distribution, thrust augmentation and noise suppression characteristics of four simple, multi-element, jet mixer-ejector configurations. The four configurations included the effect of ejector area ratio (AR = ejector area/primary jet area) and the effect of non-parallel ejector walls. We also studied in detail the configuration that produced the best noise suppression characteristics. Our results show that ejector configurations that produced the maximum maximum pumping (entrained flow per secondary inlet area) also exhibited the lowest wall pressures in the inlet region, and the maximum thrust augmentation. When cases having the same total mass flow were compared, we found that noise suppression trends corresponded with those for pumping. Surprisingly, the mixing (quantified by the peak Mach number, and flow uniformity) at the ejector exit exhibited no relationship to the noise suppression at moderate primary jet fully expanded Mach numbers (Mj is less than 1.4). However, the noise suppression dependence on the mixing was apparent at higher Mj. The above observations are justified by noting that the mixing at the ejector exit is ot a strong factor in determining the radiated noise when noise produced internal to the ejector dominates the noise field outside the ejector.

Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

The present NASA GRC-funded three-year research project is focused on studying PDE driven ejectors applicable to a hybrid Pulse Detonation/Turbofan Engine. The objective of the study is to characterize the PDE-ejector thrust augmentation. A PDE-ejector system has been designed to provide critical experimental data for assessing the performance enhancements possible with this technology. Completed tasks include demonstration of a thrust stand for measuring average thrust for detonation tube multi-cycle operation, and design of a 72-in.-long, 2.25-in.-diameter (ID) detonation tube and modular ejector assembly. This assembly will allow testing of both straight and contoured ejector geometries. Initial ejectors that have been fabricated are 72-in.-long-constant-diameter tubes (4-, 5-, and 6-in.-diameter) instrumented with high-frequency pressure transducers. The assembly has been designed such that the detonation tube exit can be positioned at various locations within the ejector tube. PDE-ejector system experiments with gaseous ethylene/ nitrogen/oxygen propellants will commence in the very near future. The program benefits from collaborations with Prof. Merkle of University of Tennessee whose PDE-ejector analysis helps guide the experiments. The present research effort will increase the TRL of PDE-ejectors from its current level of 2 to a level of 3.

Parametric Investigation of Thrust Augmentation by Ejectors on a Pulsed Detonation Tube

NASA Technical Reports Server (NTRS)

Wilson, Jack; Sgondea, Alexandru; Paxson, Daniel E.; Rosenthal, Bruce N.

2005-01-01

A parametric investigation has been made of thrust augmentation of a 1 inch diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentations for each ejector were fitted using a polynomial response surface, from which the optimum ejector diameters, and nose radius, were found. Thrust augmentations above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector.

GRC RBCC Concept Multidisciplinary Analysis

NASA Technical Reports Server (NTRS)

Suresh, Ambady

2001-01-01

This report outlines the GRC RBCC Concept for Multidisciplinary Analysis. The multidisciplinary coupling procedure is presented, along with technique validations and axisymmetric multidisciplinary inlet and structural results. The NPSS (Numerical Propulsion System Simulation) test bed developments and code parallelization are also presented. These include milestones and accomplishments, a discussion of running R4 fan application on the PII cluster as compared to other platforms, and the National Combustor Code speedup.

DTNSRDC Library Subject Thesaurus.

DTIC Science & Technology

1980-03-01

COLLISIONS BISTATIC SONAR BT SONAR BIRDS BT VERTEBRATES BITES AND STINGS NT CHICKENS BT WOUNDS AND INJURIES DUCKS GEESE BITUMENS GULLS ST ORGANIC...NT DECAPODA EL SALVADOR GUATEMALA CEPSTRUM TECHNIQUE HONDURAS MEXICO CERAMIC BODIES NICARAGUA RT BODIES PANAMA CERAMIC BONDING CENTRAL EUROPE BT...EJECTORS E EJECTORS(ORDNANCE) EGG ALBUMIN BT ALBUMINS EJECTORS(DRONANCE) BT EJECTORS EGG FO0 YOUNG NT BOMB EJECTORS USE EGGS( CHICKEN ) AND COO

Recent developments in ejector technology in the Air Force: An overview

NASA Technical Reports Server (NTRS)

Nagaraja, K. S.

1979-01-01

Basic and applied studies in thrust augmentation conducted at the Aerospace Research Laboratory at Wright-Patterson AFB which led to an effective configuration of the jet flap diffuser ejector, are reviewed. A method for compressible ejector flow analysis, developed in support of the preliminary design of an ejector thrust aircraft, is discussed and applied to single- and two-stage ejectors.

Design of an air ejector for boundary-layer bleed of an acoustically treated turbofan engine inlet during ground testing

NASA Technical Reports Server (NTRS)

Stakolich, E. G.

1978-01-01

An air ejector was designed and built to remove the boundary-layer air from the inlet a turbofan engine during an acoustic ground test program. This report describes; (1) how the ejector was sized; (2) how the ejector performed; and (3) the performance of a scale model ejector built and tested to verify the design. With proper acoustic insulation, the ejector was effective in reducing boundary layer thickness in the inlet of the turbofan engine while obtaining the desired acoustic test conditions.

Analysis of a domestic refrigerator cycle with an ejector

DOE Office of Scientific and Technical Information (OSTI.GOV)

Tomasek, M.L.; Radermacher, R.

1995-08-01

In this paper, an improved cooling cycle for a conventional domestic refrigerator-freezer utilizing an ejector for vapor precompression is analyzed using an idealized model Its energy efficiency is compared to that of the conventional refrigerator-freezer system. Emphasis is placed on off-design conditions. The ejector-enhanced refrigeration cycle consists of two evaporators that operate at different pressure and temperature levels. The ejector combines the vapor flows exiting the two evaporators into one at an intermediate pressure level The ejector cycle gives an increase of up to 12.4% in the coefficient of performance (COP) compared to that of a standard refrigerator-freezer refrigeration cycle.more » The analysis includes calculations on the optimum throat diameters of the ejector. The investigation on the off-design performance of the ejector cycle shows little dependency of energy consumption on constant ejector throat diameters.« less

Internal performance characteristics of vectored axisymmetric ejector nozzles

NASA Technical Reports Server (NTRS)

Lamb, Milton

1993-01-01

A series of vectoring axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at NASA-Langley Research Center. These ejector nozzles used convergent-divergent nozzles as the primary nozzles. The model geometric variables investigated were primary nozzle throat area, primary nozzle expansion ratio, effective ejector expansion ratio (ratio of shroud exit area to primary nozzle throat area), ratio of minimum ejector area to primary nozzle throat area, ratio of ejector upper slot height to lower slot height (measured on the vertical centerline), and thrust vector angle. The primary nozzle pressure ratio was varied from 2.0 to 10.0 depending upon primary nozzle throat area. The corrected ejector-to-primary nozzle weight-flow ratio was varied from 0 (no secondary flow) to approximately 0.21 (21 percent of primary weight-flow rate) depending on ejector nozzle configuration. In addition to the internal performance and pumping characteristics, static pressures were obtained on the shroud walls.

Efficiency Enhancement of Chiller and Heat Pump Using Natural Working Fluids with Two-phase Flow Ejector

NASA Astrophysics Data System (ADS)

Yoshikawa, Choiku; Hattori, Kazuhiro; Jeong, Jongsoo; Saito, Kiyoshi; Kawai, Sunao

An ejector can transform the expansion energy of the driving flow into the pressure build-up energy of the suction flow. Therefore, by utilizing the ejector instead of the expansion valve for the vapor compression cycle, the performance of the cycle can be greatly improved. Until now, the performance of the vapor compression cycle with the ejector has not been examined sufficiently. Therefore, this paper constructs the simulation model of the vapor compression cycle with the ejector and investigates the performance of that cycle by the simulation. Working fluids are ammonia and CO2. As a result, in case of the ejector efficiency 90%, COP of the vapor compression cycle using ammonia with the ejector is 5% higher than that of the conventional cycle and COP using CO2 with the ejector is 22% higher than that of the conventional cycle.

Internal-Performance Evaluation of Two Fixed-Divergent-Shroud Ejectors

NASA Technical Reports Server (NTRS)

Mihaloew, James R.

1960-01-01

Ejectors designed for use in a Mach 2.2 aircraft were evaluated over a range of representative primary pressure ratios and ejector corrected weight-flow ratios. Basic thrust and pumping characteristics are discussed in terms of an assumed engine operating schedule to illustrate the variation of performance with Mach number. The two designs differed about 16 percent in the shroud longitudinal spacing ratio. For corrected ejector weight-flow ratios up to 0.10, the performance of the fixed-shroud ejector designs is comparable with that of a similar continuously variable ejector except at conditions corresponding to acceleration with afterburning from Mach 0.4 to 1.2. In this region, the ejector thrust ratio decreased to a minimum of 0.96.

Viscid/inviscid interaction analysis of thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Bevilacqua, P. M.; Dejoode, A. D.

1979-01-01

A method was developed for calculating the static performance of thrust augmenting ejectors by matching a viscous solution for the flow through the ejector to an inviscid solution for the flow outside the ejector. A two dimensional analysis utilizing a turbulence kinetic energy model is used to calculate the rate of entrainment by the jets. Vortex panel methods are then used with the requirement that the ejector shroud must be a streamline of the flow induced by the jets to determine the strength of circulation generated around the shroud. In effect, the ejector shroud is considered to be flying in the velocity field of the jets. The solution is converged by iterating between the rate of entrainment and the strength of the circulation. This approach offers the advantage of including external influences on the flow through the ejector. Comparisons with data are presented for an ejector having a single central nozzle and Coanda jet on the walls. The accuracy of the matched solution is found to be especially sensitive to the jet flap effect of the flow just downstream of the ejector exit.

Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

2005-01-01

Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study and operated at frequencies up to 50 Hz. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results at each desired frequency agree with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various ejector lengths, the radius of curvature for the ejector inlets and various detonation tube/ejector tube overlap distances. For the studied experimental matrix, the results showed a maximum thrust augmentation of 106% at an operational frequency of 30 Hz. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Ejector device for direct injection fuel jet

DOEpatents

Upatnieks, Ansis [Livermore, CA

2006-05-30

Disclosed is a device for increasing entrainment and mixing in an air/fuel zone of a direct fuel injection system. The device comprises an ejector nozzle in the form of an inverted funnel whose central axis is aligned along the central axis of a fuel injector jet and whose narrow end is placed just above the jet outlet. It is found that effective ejector performance is achieved when the ejector geometry is adjusted such that it comprises a funnel whose interior surface diverges about 7.degree. to about 9.degree. away from the funnel central axis, wherein the funnel inlet diameter is about 2 to about 3 times the diameter of the injected fuel plume as the fuel plume reaches the ejector inlet, and wherein the funnel length equal to about 1 to about 4 times the ejector inlet diameter. Moreover, the ejector is most effectively disposed at a separation distance away from the fuel jet equal to about 1 to about 2 time the ejector inlet diameter.

Internal-external flow integration for a thin ejector-flapped wing section

NASA Technical Reports Server (NTRS)

Woolard, H. W.

1979-01-01

Thin airfoil theories of an ejector flapped wing section are reviewed. The global matching of the external airfoil flow with the ejector internal flow and the overall ejector flapped wing section aerodynamic performance are examined. Mathematical models of the external and internal flows are presented. The delineation of the suction flow coefficient characteristics are discussed. The idealized lift performance of an ejector flapped wing relative to a jet augmented flapped wing are compared.

Experimental Study of a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

2005-01-01

Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen (O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results are in excellent agreement with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. Further increases are possible by tailoring the ejector geometry based on CFD predictions conducted elsewhere. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Theoretical and experimental investigation of turbulent mixing on ejector configuration and performance in a solar-driven organic-vapor ejector cycle chiller

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kucha, E.I.

1984-01-01

A general method was developed to calculate two dimensional (axisymmetric) mixing of a compressible jet in a variable cross-sectional area mixing channel of the ejector. The analysis considers mixing of the primary and secondary fluids at constant pressure and incorporates finite difference approximations to the conservation equations. The flow model is based on the mixing length approximations. A detailed study and modeling of the flow phenomenon determines the best (optimum) mixing channel geometry of the ejector. The detailed ejector performance characteristics are predicted by incorporating the flow model into a solar-powered ejector cycle cooling system computer model. Freon-11 is usedmore » as both the primary and secondary fluids. Performance evaluation of the cooling system is examined for its coefficient of performance (COP) under a variety of operating conditions. A study is also conducted on a modified ejector cycle in which a secondary pump is introduced at the exit of the evaporator. Results show a significant improvement in the overall performance over that of the conventional ejector cycle (without a secondary pump). Comparison between one and two-dimensional analyses indicates that the two-dimensional ejector fluid flow analysis predicts a better overall system performance. This is true for both the conventional and modified ejector cycles.« less

Fluid channeling system

NASA Technical Reports Server (NTRS)

Davis, Donald Y. (Inventor); Hitch, Bradley D. (Inventor)

1994-01-01

A fluid channeling system includes a fluid ejector, a heat exchanger, and a fluid pump disposed in series flow communication The ejector includes a primary inlet for receiving a primary fluid, and a secondary inlet for receiving a secondary fluid which is mixed with the primary fluid and discharged therefrom as ejector discharge. Heat is removed from the ejector discharge in the heat exchanger, and the heat exchanger discharge is compressed in the fluid pump and channeled to the ejector secondary inlet as the secondary fluid In an exemplary embodiment, the temperature of the primary fluid is greater than the maximum operating temperature of a fluid motor powering the fluid pump using a portion of the ejector discharge, with the secondary fluid being mixed with the primary fluid so that the ejector discharge temperature is equal to about the maximum operating temperature of the fluid motor.

Entrainment and thrust augmentation in pulsatile ejector flows

NASA Technical Reports Server (NTRS)

Sarohia, V.; Bernal, L.; Bui, T.

1981-01-01

This study comprised direct thrust measurements, flow visualization by use of a spark shadowgraph technique, and mean and fluctuating velocity measurements with a pitot tube and linearized constant temperature hot-wire anemometry respectively. A gain in thrust of as much as 10 to 15% was observed for the pulsatile ejector flow as compared to the steady flow configuration. From the velocity profile measurements, it is concluded that this enhanced augmentation for pulsatile flow as compared to a nonpulsatile one was accomplished by a corresponding increased entrainment by the primary jet flow. It is also concluded that the augmentation and total entrainment by a constant area ejector critically depends upon the inlet geometry of the ejector. Experiments were performed to evaluate the influence of primary jet to ejector area ratio, ejector length, and presence of a diffuser on pulsatile ejector performance.

Starting Vortex Identified as Key to Unsteady Ejector Performance

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.

2004-01-01

Unsteady ejectors are currently under investigation for use in some pulse-detonation-engine-based propulsion systems. Experimental measurements made in the past, and recently at the NASA Glenn Research Center, have demonstrated that thrust augmentation can be enhanced considerably when the driver is unsteady. In ejector systems, thrust augmentation is defined as = T(sup Total)/T(sup j), where T(sup Total) is the total thrust of the combined ejector and driving jet and T(sup j) is the thrust due to the driving jet alone. There are three images in this figure, one for each of the named thrust sources. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left, and the shape and location of each driver is shown on the far right of each image. The emitted vortex is a clearly defined "doughnut" of highly vortical (spinning) flow. In these planar images, the vortex appears as two distorted circles, one above, and one below the axis of symmetry. Because they are spinning in the opposite direction, the two circles have vorticity of opposite sign and thus are different colors. There is also a rectangle shown in each image. Its width represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just "captures" the vortex: that is, the diameter bounding the outermost edge of the vortex structure. The exact mechanism behind the enhanced performance is unclear; however, it is believed to be related to the powerful vortex emitted with each pulse of the unsteady driver. As such, particle imaging velocimetry (PIV) measurements were obtained for three unsteady drivers: a pulsejet, a resonance tube, and a speaker-driven jet. All the drivers were tested with ejectors, and all exhibited performance enhancement over similarly sized steady drivers. The characteristic starting vortices of each driver are shown in these images. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left. The shape and location of each driver is shown on the far right of each image. The rectangle shown in each image represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just "captures" the vortex: that is, the diameter bounding the outermost edge of the vortex structure. Although not shown, it was observed that the emitted vortex spread as it traveled downstream. The spreading rate for the pulsejet is shown as the dashed lines in the top image. A tapered ejector was fabricated that matched this shape. When tested, the ejector demonstrated superior performance to all those previously tested at Glenn (which were essentially of straight, cylindrical form), achieving a remarkable thrust augmentation of 2. The measured thrust augmentation is shown as a function of ejector length. Also shown are the thrust augmentation values achieved with the straight, cylindrical ejectors of varying diameters. Here, thrust augmentation is plotted as a function of ejector length for several families of ejector diameters. It can be seen that large thrust augmentation values are indeed obtained and that they are sensitive to both ejector length and diameter, particularly the latter. Five curves are shown. Four correspond to straight ejector diameters of 2.2, 3.0, 4.0, and 6.0 in. The fifth curve corresponds to the tapered ejector contoured to bound the emitted vortex. For each curve, there are several data points corresponding to different lengths. The largest value of thrust augmentation is 2.0 for the tapered ejector and 1.81 for the straight ejectors. Regardless of their diameters, all the ejectors trend toward peak performance at a particular leng. That the cross-sectional dimensions of optimal ejectors scaled precisely with the vortex dimensions on three separate pulsed thrust sources demonstrates that the action of the vortex is responsible for the enhanced ejector performance. The result also suggests that, in the absence of a complete understanding of the entrainment and augmentation mechanisms, methods of characterizing starting vortices may be useful for correlating and predicting unsteady ejector performance.

A cold ejector for closed-cycle helium refrigerators

NASA Technical Reports Server (NTRS)

Johnson, D. L.; Daggett, D. L.

1987-01-01

The test results are presented of an initial cold helium ejector design that can be installed on a closed cycle refrigerator to provide refrigeration at temperatures below 4.2 K. The ejector, test apparatus, instrumentation, and test results are described. Tests were conducted both at room temperature and at cryogenic temperatures to provide operational experience with the ejector as well as for future use in the subsequent design of an ejector that will provide refrigeration at temperatures below 3 K.

Design and test of a prototype scale ejector wing

NASA Technical Reports Server (NTRS)

Mefferd, L. A.; Alden, R. E.; Bevilacqua, P. M.

1979-01-01

A two dimensional momentum integral analysis was used to examine the effect of changing inlet area ratio, diffuser area ratio, and the ratio of ejector length to width. A relatively wide range of these parameters was considered. It was found that for constant inlet area ratio the augmentation increases with the ejector length, and for constant length: width ratio the augmentation increases with inlet area ratio. Scale model tests were used to verify these trends and to examine th effect of aspect ratio. On the basis of these results, an ejector configuration was selected for fabrication and testing at a scale representative of an ejector wing aircraft. The test ejector was powered by a Pratt-Whitney F401 engine developing approximately 12,000 pounds of thrust. The results of preliminary tests indicate that the ejector develops a thrust augmentation ratio better than 1.65.

Experiments on high speed ejectors

NASA Technical Reports Server (NTRS)

Wu, J. J.

1986-01-01

Experimental studies were conducted to investigate the flow and the performance of thrust augmenting ejectors for flight Mach numbers in the range of 0.5 to 0.8, primary air stagnation pressures up to 107 psig (738 kPa), and primary air stagnation temperatures up to 1250 F (677 C). The experiment verified the existence of the second solution ejector flow, where the flow after complete mixing is supersonic. Thrust augmentation in excess of 1.2 was demonstrated for both hot and cold primary jets. The experimental ejector performed better than the corresponding theoretical optimal first solution ejector, where the mixed flow is subsonic. Further studies are required to realize the full potential of the second solution ejector. The research program was started by the Flight Dynamics Research Corporation (FDRC) to investigate the characteristic of a high speed ejector which augments thrust of a jet at high flight speeds.

Performance Enhancement of Unsteady Ejectors Investigated Using a Pulsejet Driver

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.

2003-01-01

Unsteady ejectors are currently under investigation for use in some pulse detonation engine (PDE) propulsion systems. This is due primarily to their potential high performance in comparison to steady ejectors of similar dimensions relative to the source or driver jet. Although some experimental work has been done in the past to study thrust augmentation with unsteady ejectors, there is no proven theory by which optimal design parameters can be selected and an effective ejector constructed for a given pulsed flow. Therefore, an experimental facility was developed at the NASA Glenn Research Center to study the correlation between ejector design and performance, and to get a better understanding of the flow phenomena that result in thrust augmentation. A commercially available pulsejet was used for the unsteady driving jet. This was paired with a basic, yet flexible, ejector design that allowed parametric evaluation of the effects that length, diameter, and inlet radius have on performance.

An experimental investigation of two-dimensional thrust augmenting ejectors, part 2

NASA Technical Reports Server (NTRS)

Bernal, L.; Sarohia, V.

1984-01-01

The flow-field within a two-dimensional thrust augmenting ejector has been documented experimentally. Results are presented on the mean velocity field and the turbulent correlations by Laser Doppler Velocimeter, surface pressure distribution, surface temperature distribution, and thrust performance for two shroud geometries. The maximum primary nozzle pressure ratio tested was 3.0. The tests were conducted at primary nozzle temperature ratios of 1.0, 1.8 and 2.7. Two ejector characteristic lengths have been identified based on the dynamics of the ejector flow field, i.e., a minimum length L sub m below which no significant mixing occurs, and a critical length L sub c associated with the development of U'V' correlation in the ejector. These characteristic lengths divide the ejector flow field into three distinctive regions: the entrance region where there is no direct interaction between the primary flow and the ejector shroud; the interaction region where there is an increased momentum of induced flow near the shroud surface; and a pipe flow region characterized by an increased skin friction where x is the distance downstream from the ejector inlet. The effect of the coflowing induced flow has been shown to produce inside the ejector a centerline velocity that has increased over the free-jet data.

Pulsed Ejector Thrust Amplification Tested and Modeled

NASA Technical Reports Server (NTRS)

Wilson, Jack

2004-01-01

There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter.

Ejectors of power plants turbine units efficiency and reliability increasing

NASA Astrophysics Data System (ADS)

Aronson, K. E.; Ryabchikov, A. Yu.; Kuptsov, V. K.; Murmanskii, I. B.; Brodov, Yu. M.; Zhelonkin, N. V.; Khaet, S. I.

2017-11-01

The functioning of steam turbines condensation systems influence on the efficiency and reliability of a power plant a lot. At the same time, the condensation system operating is provided by basic ejectors, which maintain the vacuum level in the condenser. Development of methods of efficiency and reliability increasing for ejector functioning is an actual problem of up-to-date power engineering. In the paper there is presented statistical analysis of ejector breakdowns, revealed during repairing processes, the influence of such damages on the steam turbine operating reliability. It is determined, that 3% of steam turbine equipment breakdowns are the ejector breakdowns. At the same time, about 7% of turbine breakdowns are caused by different ejector malfunctions. Developed and approved design solutions, which can increase the ejector functioning indexes, are presented. Intercoolers are designed in separated cases, so the air-steam mixture can’t move from the high-pressure zones to the low-pressure zones and the maintainability of the apparatuses is increased. By U-type tubes application, the thermal expansion effect of intercooler tubes is compensated and the heat-transfer area is increased. By the applied nozzle fixing construction, it is possible to change the distance between a nozzle and a mixing chamber (nozzle exit position) for operating performance optimization. In operating conditions there are provided experimental researches of more than 30 serial ejectors and also high-efficient 3-staged ejector EPO-3-80, designed by authors. The measurement scheme of the designed ejector includes 21 indicator. The results of experimental tests with different nozzle exit positions of the ejector EPO-3-80 stream devices are presented. The pressure of primary stream (water steam) is optimized. Experimental data are well-approved by the calculation results.

Design and experimental investigation of an ejector in an air-conditioning and refrigeration system

DOE Office of Scientific and Technical Information (OSTI.GOV)

AL-Khalidy, N.; Zayonia, A.

1995-12-31

This paper discusses the conservation of energy in a refrigerant ejector refrigerating machine using heat driven from the concentrator collectors. The working refrigerant was R-113. The design of an ejector operating in an air-conditioning and refrigerating system with a low thermal source (70 C to 100 C) is presented. The influence of three major parameters--boiler, condenser, and evaporator temperature--on ejector efficiency is discussed. Experimental results show that the condenser temperature is the major influence at a low evaporator temperature. The maximum ejector efficiency was 31%.

A model for prediction of STOVL ejector dynamics

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1989-01-01

A semi-empirical control-volume approach to ejector modeling for transient performance prediction is presented. This new approach is motivated by the need for a predictive real-time ejector sub-system simulation for Short Take-Off Verticle Landing (STOVL) integrated flight and propulsion controls design applications. Emphasis is placed on discussion of the approximate characterization of the mixing process central to thrust augmenting ejector operation. The proposed ejector model suggests transient flow predictions are possible with a model based on steady-flow data. A practical test case is presented to illustrate model calibration.

An experimental study of a three-dimensional thrust augmenting ejector using laser Doppler velocimetry

NASA Technical Reports Server (NTRS)

Storms, Bruce Lowell

1989-01-01

Flow field measurements were obtained in a three-dimensional thrust augmenting ejector using laser Doppler velocimetry and hot wire anemometry. The primary nozzle, segmented into twelve slots of aspect ratio 3.0, was tested at a pressure ratio of 1.15. Results are presented on the mean velocity, turbulence intensity, and Reynolds stress progressions in the mixing chamber of the constant area ejector. The segmented nozzle was found to produce streamwise vortices that may increase the mixing efficiency of the ejector flow field. Compared to free jet results, the jet development is reduced by the presence of the ejector walls. The resulting thrust augmentation ratio of this ejector was also calculated to be 1.34.

Highly Variable Cycle Exhaust Model Test (HVC10)

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Wernet, Mark; Podboy, Gary; Bozak, Rick

2010-01-01

Results from acoustic and flow-field studies using the Highly Variable Cycle Exhaust (HVC) model were presented. The model consisted of a lobed mixer on the core stream, an elliptic nozzle on the fan stream, and an ejector. For baseline comparisons, the fan nozzle was replaced with a round nozzle and the ejector doors were removed from the model. Acoustic studies showed far-field noise levels were higher for the HVC model with the ejector than for the baseline configuration. Results from Particle Image Velocimetry (PIV) studies indicated that large flow separation regions occurred along the ejector doors, thus restricting flow through the ejector. Phased array measurements showed noise sources located near the ejector doors for operating conditions where tones were present in the acoustic spectra.

Dual mode nuclear rocket system applications.

NASA Technical Reports Server (NTRS)

Boretz, J. E.; Bell, J. M.; Plebuch, R. K.; Priest, C. C.

1972-01-01

Mission areas where the dual-mode nuclear rocket system is superior to nondual-mode systems are demonstrated. It is shown that the dual-mode system is competitive with the nondual-mode system even for those specific missions and particular payload configurations where it does not have a clear-cut advantage.

Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Pal, S.; Santoro, Robert J.; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

2005-01-01

Thrust augmentation results of an ongoing study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) setup with various ejector configurations. The PDE used in these experiments utilizes ethylene (C2H4) as the fuel, and an equi-molar mixture of oxygen and nitrogen as the oxidizer at an equivalence ratio of one. High fidelity thrust measurements were made using an integrated spring damper system. The baseline thrust of the PDE engine was first measured and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The parameter space for the study included ejector length, PDE tube exit to ejector tube inlet overlap distance, and straight versus rounded ejector inlets. The relationship between the thrust augmentation results and various physical phenomena is described. To further understand the flow dynamics, shadow graph images of the exiting shock wave front from the PDE were also made. For the studied parameter space, the results showed a maximum augmentation of 40%. Further increase in augmentation is possible if the geometry of the ejector is tailored, a topic currently studied by numerous groups in the field.

Unsteady Ejector Performance: an Experimental Investigation Using a Pulsejet Driver

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wilson, Jack; Dougherty, Kevin T.

2002-01-01

An experimental investigation is described in which thrust augmentation and mass entrainment were measured for a variety of simple cylindrical ejectors driven by a gasoline-fueled pulsejet. The ejectors were of varying length, diameter, and inlet radius. Measurements were also taken to determine the effect on performance of the distance between pulsejet exit and ejector inlet. Limited tests were also conducted to determine the effect of driver cross-sectional shape. Optimal values were found for all three ejector parameters with respect to thrust augmentation. This was not the case with mass entrainment, which increased monotonically with ejector diameter. Thus, it was found that thrust augmentation is not necessarily directly related to mass entrainment, as is often supposed for ejectors. Peak thrust augmentation values of 1.8 were obtained. Peak mass entrainment values of 30 times the driver mass flow were also observed. Details of the experimental setup and results are presented. Preliminary analysis of the results indicates that the enhanced performance obtained with an unsteady jet (primary source) over comparably sized ejectors driven with steady jets is due primarily to the structure of the starting vortex-type flow associated with the former.

Development of acoustically lined ejector technology for multitube jet noise suppressor nozzles by model and engine tests over a wide range of jet pressure ratios and temperatures

NASA Technical Reports Server (NTRS)

Atvars, J.; Paynter, G. C.; Walker, D. Q.; Wintermeyer, C. F.

1974-01-01

An experimental program comprising model nozzle and full-scale engine tests was undertaken to acquire parametric data for acoustically lined ejectors applied to primary jet noise suppression. Ejector lining design technology and acoustical scaling of lined ejector configurations were the major objectives. Ground static tests were run with a J-75 turbojet engine fitted with a 37-tube, area ratio 3.3 suppressor nozzle and two lengths of ejector shroud (L/D = 1 and 2). Seven ejector lining configurations were tested over the engine pressure ratio range of 1.40 to 2.40 with corresponding jet velocities between 305 and 610 M/sec. One-fourth scale model nozzles were tested over a pressure ratio range of 1.40 to 4.0 with jet total temperatures between ambient and 1088 K. Scaling of multielement nozzle ejector configurations was also studied using a single element of the nozzle array with identical ejector lengths and lining materials. Acoustic far field and near field data together with nozzle thrust performance and jet aerodynamic flow profiles are presented.

Characteristics of Five Ejector Configurations at Free-Stream Mach Numbers from 0 to 2.0

NASA Technical Reports Server (NTRS)

Klann, John L.; Huff, Ronald G.

1959-01-01

Thrust, air-handling, and base-pressure characteristics of five ejector configurations were investigated in the Lewis 8-by 6-foot wind tunnel at free-stream Mach numbers from 0 to 2.0 over ranges of primary-jet pressure ratio up to 24 and corrected secondary weight-flow ratio up to 13 percent. The ejector-shroud geometries varied from convergent to divergent. Base pressure ratio and ejector performance were interrelated by means of an exit-momentum parameter. Correlations, to at least a first approximation, with base pressure ratio, of both internal-ejector-flow separation and external-flow separation over the model boattail were shown. Furthermore, it was shown that magnitudes and exact trends in base pressure ratio depended largely, and in a complicated fashion, on ejector geometry and amount of secondary flow. External-stream effects on ejector jet thrust were determined for a typical schedule of jet-engine pressure ratios. With the exception of the ejector having the largest (1.81) shroud-exit-to primary-diameter ratio, there were no stream effects at Mach numbers from 1.5 to 2.0 and variations from quiescent-air thrust data were less than 2.5 percent at the subsonic speed investigated.

Internal performance characteristics of thrust-vectored axisymmetric ejector nozzles

NASA Technical Reports Server (NTRS)

Lamb, Milton

1995-01-01

A series of thrust-vectored axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at the Langley research center. This study indicated that discontinuities in the performance occurred at low primary nozzle pressure ratios and that these discontinuities were mitigated by decreasing expansion area ratio. The addition of secondary flow increased the performance of the nozzles. The mid-to-high range of secondary flow provided the most overall improvements, and the greatest improvements were seen for the largest ejector area ratio. Thrust vectoring the ejector nozzles caused a reduction in performance and discharge coefficient. With or without secondary flow, the vectored ejector nozzles produced thrust vector angles that were equivalent to or greater than the geometric turning angle. With or without secondary flow, spacing ratio (ejector passage symmetry) had little effect on performance (gross thrust ratio), discharge coefficient, or thrust vector angle. For the unvectored ejectors, a small amount of secondary flow was sufficient to reduce the pressure levels on the shroud to provide cooling, but for the vectored ejector nozzles, a larger amount of secondary air was required to reduce the pressure levels to provide cooling.

Electromechanical Dynamics Simulations of Superconducting LSM Rocket Launcher System in Attractive-Mode

NASA Technical Reports Server (NTRS)

Yoshida, Kinjiro; Hayashi, Kengo; Takami, Hiroshi

1996-01-01

Further feasibility study on a superconducting linear synchronous motor (LSM) rocket launcher system is presented on the basis of dynamic simulations of electric power, efficiency and power factor as well as the ascending motions of the launcher and rocket. The advantages of attractive-mode operation are found from comparison with repulsive-mode operation. It is made clear that the LSM rocket launcher system, of which the long-stator is divided optimally into 60 sections according to launcher speeds, can obtain high efficiency and power factor.

Investigation of the gas-jet ejector in KamAZ trucks

DOE Office of Scientific and Technical Information (OSTI.GOV)

Shkret, L.Y.; Berezea, A.I.; Lobkov, A.N.

1984-03-01

This article considers the possibility of using gas-jet vacuum pumps in tank trucks for transporting liquids (water) at drilling sites. The discharge system of the KamAZ trucks can be reliably sealed by an engine brake, an important prerequisite of reliable operation of a gas-jet ejector that is switched on when the tank is being filled. The ejector consists of a housing, a Laval nozzle, a front wall with cylindrical neck, a tin-plate diffuser, an air supply pipe, and a flange for attaching the ejector to the flange of the exhaust muffler of the truck. The gas-jet ejectors are driven bymore » the exhaust gas (EG) of the trucks. The dependences of the EG flow rate, fuel expenditure, EG temperature ahead of the ejector, and the rotational frequency of the engine crankshaft on the diameter at different EG pressures. It is recommended that gas-jet ejectors be used on series produced tank trucks instead of rotary vacuum pumps with mechanical drive.« less

Effect of Operating Frequency and Fill Time on PDE-Ejector Thrust Performance

NASA Technical Reports Server (NTRS)

Landry, K.; Santoro, Robert J.; Pal, Sibtosh; Shehadeh, R.; Bouvet, N.; Lee, S.-Y.

2005-01-01

Thrust measurements for a pulse detonation engine (PDE)-ejector system were determined for a range of operating frequencies. Various length tubular ejectors were utilized. The results were compared to the measurements of the thrust output of the PDE alone to determine the enhancement provided by each ejector configuration at the specified frequencies. Ethylene was chosen as the fuel, with an equi-molar mixture of nitrogen and oxygen acting as the oxidizer. The propellant was kept at an equivalence ratio of one during all the experiments. The system was operated for frequencies between 20 and 50 Hz. The parameter space of the study included PDE operation frequency, ejector length, overlap percentage, the radius of curvature for the ejector inlets, and duration of the time allowed between cycles. The results of the experiments showed a maximum thrust augmentation of 120% for a PDE-ejector configuration at a frequency of 40Hz with a fill time of 10 ms.

A computational study of thrust augmenting ejectors based on a viscous-inviscid approach

NASA Technical Reports Server (NTRS)

Lund, Thomas S.; Tavella, Domingo A.; Roberts, Leonard

1987-01-01

A viscous-inviscid interaction technique is advocated as both an efficient and accurate means of predicting the performance of two-dimensional thrust augmenting ejectors. The flow field is subdivided into a viscous region that contains the turbulent jet and an inviscid region that contains the ambient fluid drawn into the device. The inviscid region is computed with a higher-order panel method, while an integral method is used for the description of the viscous part. The strong viscous-inviscid interaction present within the ejector is simulated in an iterative process where the two regions influence each other en route to a converged solution. The model is applied to a variety of parametric and optimization studies involving ejectors having either one or two primary jets. The effects of nozzle placement, inlet and diffuser shape, free stream speed, and ejector length are investigated. The inlet shape for single jet ejectors is optimized for various free stream speeds and Reynolds numbers. Optimal nozzle tilt and location are identified for various dual-ejector configurations.

Aerodynamic and acoustic performance of ejectors for engine-under-the-wing concepts

NASA Technical Reports Server (NTRS)

Vonglahn, U.; Goodykoontz, J. H.; Groesbeck, D.

1974-01-01

Subsonic thrust augmentation, exhaust plume velocity contours and acoustic characteristics of a small-scale, 6-tube mixer nozzle with ejector were obtained with and without a wing. Thrust augmentation up to 30 percent was achieved. Aerodynamic results showed that at a given location, greater downstream velocities are obtained with an ejector than with the baseline nozzle. Ejectors reduce high frequency noise; however, low frequency noise amplification also occurs. Acoustic reflections off the wing increase the noise level to a ground observer. With an ejector, the acoustic benefits of forward velocity may be significantly reduced compared with the baseline nozzle.

Improvements to the ejector expansion refrigeration cycle

DOE Office of Scientific and Technical Information (OSTI.GOV)

Menegay, P.; Kornhauser, A.A.

1996-12-31

The ejector expansion refrigeration cycle (EERC) is a variant of the standard vapor compression cycle in which an ejector is used to recover part of the work that would otherwise be lost in the expansion valve. In initial testing EERC performance was poor, mainly due to thermodynamic non-equilibrium conditions in the ejector motive nozzle. Modifications were made to correct this problem, and significant performance improvements were found.

1. EXTERIOR CONTEXT VIEW OF BUILDING 620, THE SEWAGE EJECTOR, ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

1. EXTERIOR CONTEXT VIEW OF BUILDING 620, THE SEWAGE EJECTOR, LOOKING NORTHEAST. - Mill Valley Air Force Station, Sewage Ejector, East Ridgecrest Boulevard, Mount Tamalpais, Mill Valley, Marin County, CA

Rotary wave-ejector enhanced pulse detonation engine

NASA Astrophysics Data System (ADS)

Nalim, M. R.; Izzy, Z. A.; Akbari, P.

2012-01-01

The use of a non-steady ejector based on wave rotor technology is modeled for pulse detonation engine performance improvement and for compatibility with turbomachinery components in hybrid propulsion systems. The rotary wave ejector device integrates a pulse detonation process with an efficient momentum transfer process in specially shaped channels of a single wave-rotor component. In this paper, a quasi-one-dimensional numerical model is developed to help design the basic geometry and operating parameters of the device. The unsteady combustion and flow processes are simulated and compared with a baseline PDE without ejector enhancement. A preliminary performance assessment is presented for the wave ejector configuration, considering the effect of key geometric parameters, which are selected for high specific impulse. It is shown that the rotary wave ejector concept has significant potential for thrust augmentation relative to a basic pulse detonation engine.

Hybrid Vapor Compression Ejector Cycle: Presentation to IAPG Mechanical Working Group

DTIC Science & Technology

2011-08-01

Compression Ejector Cycle: Presentation to IAPG Mechanical Working Group Parmesh Verma and Tom Radcliff, United Technologies Research Center UNCLASSIFIED... Ejector Cycle Presentation to IAPG Mechanical Working Group 5a. CONTRACT NUMBER W909MY-10-C-0005 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6...hybrid vapor compression ejector heat pump cycle developed under an American Recovery and Reinvestment Act funded contract is provided. 15. SUBJECT

Credit WCT. Photographic copy of photograph, view east showing the ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit WCT. Photographic copy of photograph, view east showing the Y-stage ejector nozzle as the Y-stage ejector is being installed in the Dd ejector train in 1962. In the distance can be seen the western end of the Z-stage ejector. (JPL negative no. 384-3345-A, 8 November 1962) - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Energetics of Vortex Ring Formation.

DTIC Science & Technology

1983-11-01

Sorohia, V., "An Experimental Investigation of Thrust Augmenting Ejector Flows", Proceedings of the Ejector Workshop for Aerospace Applications, AFWAL-TR...induction thrust augmentrs, su’h comparing thr mass and energy content of fully formed as the ejector , the migration of finite sized eddie, laminar vortex...Intermittent Jet to a Secondary Fluid in an Ejector Type Thrust Augmentor", Hiller Aircraft Company, Interim Report ARD-305, June 1962. 3. Bernal, L. and

Comparison of the initial development of shear layers in two-dimensional and axisymmetric ejector configurations

NASA Technical Reports Server (NTRS)

Dufflocq, M.; Benjamin, M. A.; Roan, V. P.

1993-01-01

A two-phase experimental investigation designed to study the development of shear layers in axisymmetric and two-dimensional single-nozzle ejectors has been completed. In this study, combinations of similar and dissimilar gases were used as the supersonic primary and subsonic secondary. Test cases included combinations of air/air, argon/air and helium/air as the supersonic primary and subsonic secondary, respectively. Similar flow conditions were studied for each ejector configuration. Mixing of the gases occurred in a constant-area tube, where the inlet pressure was maintained at 34.5 kPa. The cases studied resulted in convective Mach numbers that range between 0.06 and 1.9. The data gathered shows differences between the initial shear-layer development for the two ejector geometries, and also between the different test cases studied for each ejector configuration. The measured growth rates for the axisymmetric ejector are more than twice those measured for the two-dimensional ejector. However, in both cases the results show that compressibility has a reducing effect on the growth rate. Further, in the region immediately after the inlet to the mixing tube, compressibility seems to affect the ejector shear layers in a manner similar to that of two-stream two-dimensional mixing layers.

The Effect of Pulse Length and Ejector Radius on Unsteady Ejector Performance

NASA Technical Reports Server (NTRS)

Wilson, Jack

2005-01-01

The thrust augmentation of a set of ejectors driven by a shrouded Hartmann-Sprenger tube has been measured at four different frequencies. Each frequency corresponded to a different length to diameter ratio of the pulse of air leaving the driver shroud. Two of the frequencies had length to diameter ratios below the formation number, and two above. The formation number is the value of length to diameter ratio below which the pulse converts to a vortex ring only, and above which the pulse becomes a vortex ring plus a trailing jet. A three level, three parameter Box-Behnken statistical design of experiment scheme was performed at each frequency, measuring the thrust augmentation generated by the appropriate ejectors from the set. The three parameters were ejector length, radius, and inlet radius. The results showed that there is an optimum ejector radius and length at each frequency. Using a polynomial fit to the data, the results were interpolated to different ejector radii and pulse length to diameter ratios. This showed that a peak in thrust augmentation occurs when the pulse length to diameter ratio equals the formation number, and that the optimum ejector radius is 0.87 times the sum of the vortex ring radius and the core radius.

Ultra-high-throughput microarray generation and liquid dispensing using multiple disposable piezoelectric ejectors.

PubMed

Hsieh, Huangpin Ben; Fitch, John; White, Dave; Torres, Frank; Roy, Joy; Matusiak, Robert; Krivacic, Bob; Kowalski, Bob; Bruce, Richard; Elrod, Scott

2004-03-01

The authors have constructed an array of 12 piezoelectric ejectors for printing biological materials. A single-ejector footprint is 8 mm in diameter, standing 4 mm high with 2 reservoirs totaling 76 micro L. These ejectors have been tested by dispensing various fluids in several environmental conditions. Reliable drop ejection can be expected in both humidity-controlled and ambient environments over extended periods of time and in hot and cold room temperatures. In a prototype system, 12 ejectors are arranged in a rack, together with an X - Y stage, to allow printing any pattern desired. Printed arrays of features are created with a biological solution containing bovine serum albumin conjugated oligonucleotides, dye, and salty buffer. This ejector system is designed for the ultra-high-throughput generation of arrays on a variety of surfaces. These single or racked ejectors could be used as long-term storage vessels for materials such as small molecules, nucleic acids, proteins, or cell libraries, which would allow for efficient preprogrammed selection of individual clones and greatly reduce the chance of cross-contamination and loss due to transfer. A new generation of design ideas includes plastic injection molded ejectors that are inexpensive and disposable and handheld personal pipettes for liquid transfer in the nanoliter regime.

Preliminary dynamic tests of a flight-type ejector

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1992-01-01

A thrust augmenting ejector was tested to provide experimental data to assist in the assessment of theoretical models to predict duct and ejector fluid-dynamic characteristics. Eleven full-scale thrust augmenting ejector tests were conducted in which a rapid increase in the ejector nozzle pressure ratio was effected through a unique facility, bypass/burst-disk subsystem. The present work examines two cases representative of the test performance window. In the first case, the primary nozzle pressure ration (NPR) increased 36 percent from one unchoked (NPR = 1.29) primary flow condition to another (NPR = 1.75) over a 0.15 second interval. The second case involves choked primary flow conditions, where a 17 percent increase in primary nozzle flowrate (from NPR = 2.35 to NPR = 2.77) occurred over approximately 0.1 seconds. Although the real-time signal measurements support qualitative remarks on ejector performance, extracting quantitative ejector dynamic response was impeded by excessive aerodynamic noise and thrust stand dynamic (resonance) characteristics. It does appear, however, that a quasi-steady performance assumption is valid for this model with primary nozzle pressure increased on the order of 50 lb(sub f)/s. Transient signal treatment of the present dataset is discussed and initial interpretations of the results are compared with theoretical predictions for a similar Short Takeoff and Vertical Landing (STOVL) ejector model.

Achieving an adequate minute volume through a 2 mm transtracheal catheter in simulated upper airway obstruction using a modified industrial ejector.

PubMed

Hamaekers, A E W; Götz, T; Borg, P A J; Enk, D

2010-03-01

Needle cricothyrotomy and subsequent transtracheal jet ventilation (TTJV) is one of the last options to restore oxygenation while managing an airway emergency. However, in cases of complete upper airway obstruction, conventional TTJV is ineffective and dangerous. We transformed a small, industrial ejector into a simple, manual ventilator providing expiratory ventilation assistance (EVA). An ejector pump was modified to allow both insufflation of oxygen and jet-assisted expiration through an attached 75 mm long transtracheal catheter (TTC) with an inner diameter (ID) of 2 mm by alternately occluding and releasing the gas outlet of the ejector pump. In a lung simulator, the modified ejector pump was tested at different compliances and resistances. Inspiration and expiration times were measured and achievable minute volumes (MVs) were calculated to determine the effect of EVA. The modified ejector pump shortened the expiration time and an MV up to 6.6 litre min(-1) could be achieved through a 2 mm ID TTC in a simulated obstructed airway. The principle of ejector-based EVA seems promising and deserves further evaluation.

Steam ejector as an industrial heat pump

DOE Office of Scientific and Technical Information (OSTI.GOV)

Arnold, H.G.; Huntley, W.R.; Perez-Blanco, H.

1982-01-01

The steam ejector is analyzed for use in industrial heat recovery applications and compared to mechanical compressor heat pumps. An estimated ejector performance was analyzed using methods based on conservation of mass, momentum, and energy; using steam properties to account for continuity; and using appropriate efficiencies for the nozzle and diffuse performance within the ejector. A potential heat pump application at a paper plant in which waste water was available in a hot well downstream of the paper machine was used to describe use of the stream ejector. Both mechanical compression and jet ejector heat pumps were evaluated for recompressionmore » of flashed steam from the hot well. It is noted that another possible application of vapor recompression heat pumps is the recovery of waste heat from large facilities such as the gaseous diffusion plants. The economics of recovering waste heat in similar applications is analyzed. (MCW)« less

A new thermally driven refrigeration system with environmental benefits

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garris, C.A. Jr.; Hong, W.J.; Mavriplis, C.

1998-07-01

The pressure-exchange ejector offers the possibility of attaining a breakthrough in the level of performance of ejectors by means of utilizing non-dissipative non-steady flow mechanisms. Yet, the device retains much of the mechanical simplicity of conventional steady-flow ejectors. If such a substantial improvement in performance is demonstrated, its application to ejector refrigeration will be very important. Such a development would provide significant benefits for the environment in terms of both CFC usage reduction and greenhouse gas reduction. The current paper will discuss in detail the concept of pressure-exchange ejector refrigeration, compare it with existing technologies, and discuss the potential impactmore » that might be derived if certain levels of ejector performance can be achieved. Since the limiting issue on the system performance is in the fluid dynamics of non-steady flow induction, research issues and recent progress will be discussed.« less

Preliminary dynamic tests of a flight-type ejector

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1992-01-01

A thrust augmenting ejector was tested to provide experimental data to assist in the assessment of theoretical models to predict duct and ejector fluid-dynamic characteristics. Eleven full-scale thrust augmenting ejector tests were conducted in which a rapid increase in the ejector nozzle pressure ratio was effected through a unique bypass/burst-disk subsystem. The present work examines two cases representative of the test performance window. In the first case, the primary nozzle pressure ration (NPR) increased 36 percent from one unchoked (NPR = 1.29) primary flow condition to another (NPR = 1.75) over a 0.15 second interval. The second case involves choked primary flow conditions, where a 17 percent increase in primary nozzle flowrate (from NPR = 2.35 to NPR = 2.77) occurred over approximately 0.1 seconds. Transient signal treatment of the present dataset is discussed and initial interpretations of the results are compared with theoretical predictions for a similar STOVL ejector model.

Acoustic, Flow Related, and Performance Related Experimental Results for Generation 1.5 High Speed Civil Transport (HSCT) 2-Dimensional Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Salikuddin, M.; Wisler, S.; Majjigi, R.

2004-01-01

The principle objectives of the current program were to experimentally investigate the repeatability of acoustic and aerodynamic characteristics of 2D-CD mixer-ejector nozzles and the effects on the acoustic and aerodynamic characteristics of 2D mixer-ejectors due to (1) the configurational variations, which include mixers with aligned CD chutes, aligned convergent chutes, and staggered CD chutes and aerodynamic cycle variables, (2) treatment variations by using different treatment materials, treating the ejector with varying area, location, and treatment thickness for a mixer-ejector configuration, and (3) secondary inlet shape (i.e., a more realistic inlet) and the blockage across the inlet (a possible fin-like structure needed for installation purpose) by modifying one of the inlet of a mixer-ejector configuration. The objectives also included the measurement dynamic pressures internal to the ejector for a few selected configuration to examine the internal noise characteristics.

A Status of the Advanced Space Transportation Program from Planning to Action

NASA Technical Reports Server (NTRS)

Lyles, Garry; Griner, Carolyn

1998-01-01

A Technology Plan for Enabling Commercial Space Business was presented at the 48th International Astronautical Congress in Turin, Italy. This paper presents a status of the program's accomplishments. Technology demonstrations have progressed in each of the four elements of the program; (1) Low Cost Technology, (2) Advanced Reusable Technology, (3) Space Transfer Technology and (4) Space Transportation Research. The Low Cost Technology program element is primarily focused at reducing development and acquisition costs of aerospace hardware using a "design to cost" philosophy with robust margins, adapting commercial manufacturing processes and commercial off-the-shelf hardware. The attributes of this philosophy for small payload launch are being demonstrated at the component, sub-system, and system level. The X-34 "Fastrac" engine has progressed through major component and subsystem demonstrations. A propulsion system test bed has been implemented for system-level demonstration of component and subsystem technologies; including propellant tankage and feedlines, controls, pressurization, and engine systems. Low cost turbopump designs, commercial valves and a controller are demonstrating the potential for a ten-fold reduction in engine and propulsion system costs. The Advanced Reusable Technology program element is focused on increasing life through high strength-to-weight structures and propulsion components, highly integrated propellant tanks, automated checkout and health management and increased propulsion system performance. The validation of rocket based combined cycle (RBCC) propulsion is pro,-,ressing through component and subsystem testing. RBCC propulsion has the potential to provide performance margin over an all rocket system that could result in lower gross liftoff weight, a lower propellant mass fraction or a higher payload mass fraction. The Space Transfer Technology element of the program is pursuing technology that can improve performance and dramatically reduce the propellant and structural mass of orbit transfer and deep space systems. Flight demonstration of ion propulsion is progressing towards launch. Ion propulsion is the primary propulsion for Deep Space 1; a flyby of comet West-kohoutek-lkemura and asteroid 3352 McAuliffe. Testing of critical solar-thermal propulsion subsystems have been accomplished and planning is continuing for the flight demonstration of an electrodynamic tether orbit transfer system. The forth and final element of the program, Space Transportation Research, has progressed in several areas of propulsion research. This element of the program is focused at long-term (25 years) breakthrough concepts that could bring launch costs to a factor of one hundred below today's cost or dramatically expand planetary travel and enable interstellar travel.

An Integration of the Turbojet and Single-Throat Ramjet

NASA Technical Reports Server (NTRS)

Trefny, C. J.; Benson, T. J.

1995-01-01

A turbine-engine-based hybrid propulsion system is described. Turbojet engines are integrated with a single-throat ramjet so as to minimize variable geometry and eliminate redundant propulsion components. The result is a simple, lightweight system that is operable from takeoff to high Mach numbers. Non-afterburning turbojets are mounted within the ramjet duct. They exhaust through a converging-diverging (C-D) nozzle into a common ramjet burner section. At low speed the ejector effect of the C-D nozzle aerodynamically isolates the relatively high pressure turbojet exhaust stream from the ramjet duct. As the Mach number increases, and the turbojet pressure ratio diminishes, the system is biased naturally toward ramjet operation. The common ramjet burner is fueled with hydrogen and thermally choked, thus avoiding the weight and complexity of a variable geometry, split-flow exhaust system. The mixed-compression supersonic inlet and subsonic diffuser are also common to both the turbojet and ramjet cycles. As the compressor face total temperature limit is approached, a two-position flap within the inlet is actuated, which closes off the turbojet inlet and provides increased internal contraction for ramjet operation. Similar actuation of the turbojet C-D nozzle flap completes the enclosure of the turbojet. Performance of the hybrid system is compared herein to that of the discrete turbojet and ramjet engines from takeoff to Mach 6. The specific impulse of the hybrid system falls below that of the non-integrated turbojet and ramjet because of ejector and Rayleigh losses. Unlike the discrete turbojet or ramjet however, the hybrid system produces thrust over the entire Mach number range. An alternate mode of operation for takeoff and low speed is also described. In this mode the C-D nozzle flap is deflected to a third position, which closes off the ramjet duct and eliminates the ejector total pressure loss.

Credit BG. View west of Test Stand "D" complex, with ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. View west of Test Stand "D" complex, with ends of Dd (left) and Dy (right) station ejectors in view. Steam piping from accumulator (sphere) to ejectors is apparent; long horizontal loops in the pipes permit expansion and contraction without special joints. The small platform straddling the Dd ejector (near the accumulator) was originally constructed for a "Hyprox" steam generator which supplied steam to the Dd ejector before the accumulator and Dy stand were built. Note ejectors on top of interstage condenser in Test Stand "D" tower. Metal shed in far right background is for storage - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Augmenting ejector endwall effects. [V/STOL aircraft

NASA Technical Reports Server (NTRS)

Porter, J. L.; Squyers, R. A.

1979-01-01

Rectangular inlet ejectors which had multiple hypermixing nozzles for their primary jets were investigated for the effects of endwall blowing on thrust augmentation performance. The ejector configurations tested had both straight wall and active boundary layer control type diffusers. Endwall flows were energized and controlled by simple blowing jets suitably located in the ejector. Both the endwall and boundary layer control diffuser blowing rates were varied to determine optimum performance. High area ratio diffusers with insufficient endwall blowing showed endwall separation and rapid degradation of thrust performance. Optimized values of diffuser boundary layer control and endwall nozzle blowing rates in an ejector augmenter were shown to achieve high levels of augmentation performance for maximum compactness.

Internal Performance Evaluation of a Two Position Divergent Shroud Ejector

NASA Technical Reports Server (NTRS)

Mihaloew, James R.; Stofan, Andrew J.

1960-01-01

A two-position divergent shroud ejector was investigated in an unheated quiescent-air facility over a range of operational variables applicable to a Mach 2.5 aircraft. The performance data are shown in terms of hypothetical engine operating conditions to illustrate variations of performance with Mach number. The overall thrust performance was reasonably good, with ejector thrust ratios ranging from 0.97 to 0.98 for all conditions except that corresponding to acceleration with afterburning through the transonic flight Mach number region from 0.9 to 1.1, where the ejector thrust ratio decreased to as low as 0.945 for an ejector corrected weight-flow ratio of 0.105.

Parametric analysis of diffuser requirements for high expansion ratio space engine

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.; Anderson, P. G.

1981-01-01

A supersonic diffuser ejector design computer program was developed. Using empirically modified one dimensional flow methods the diffuser ejector geometry is specified by the code. The design code results for calculations up to the end of the diffuser second throat were verified. Diffuser requirements for sea level testing of high expansion ratio space engines were defined. The feasibility of an ejector system using two commonly available turbojet engines feeding two variable area ratio ejectors was demonstrated.

Preliminary Internal Performance Data for a Variable-Ejector Assembly on the XJ79-GE-1 Turbojet Engine. II; Afterburning Configurations

NASA Technical Reports Server (NTRS)

Bloomer, Harry E.; Groesbeck, Donald E.

1957-01-01

Internal performance of an XJ79-GE-1 variable ejector was experimentally determined with the primary nozzle in two representative after-burning positions. Jet-thrust and air-handling data were obtained in quiescent air for 4 selected ejector configurations over a wide range of secondary to primary airflow ratios and primary-nozzle pressure ratios. The experimental ejector data are presented in both graphical and tabulated form.

Parametric Studies of the Ejector Process within a Turbine-Based Combined-Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Walker, James F.; Trefny, Charles J.

1999-01-01

Performance characteristics of the ejector process within a turbine-based combined-cycle (TBCC) propulsion system are investigated using the NPARC Navier-Stokes code. The TBCC concept integrates a turbine engine with a ramjet into a single propulsion system that may efficiently operate from takeoff to high Mach number cruise. At the operating point considered, corresponding to a flight Mach number of 2.0, an ejector serves to mix flow from the ramjet duct with flow from the turbine engine. The combined flow then passes through a diffuser where it is mixed with hydrogen fuel and burned. Three sets of fully turbulent Navier-Stokes calculations are compared with predictions from a cycle code developed specifically for the TBCC propulsion system. A baseline ejector system is investigated first. The Navier-Stokes calculations indicate that the flow leaving the ejector is not completely mixed, which may adversely affect the overall system performance. Two additional sets of calculations are presented; one set that investigated a longer ejector region (to enhance mixing) and a second set which also utilized the longer ejector but replaced the no-slip surfaces of the ejector with slip (inviscid) walls in order to resolve discrepancies with the cycle code. The three sets of Navier-Stokes calculations and the TBCC cycle code predictions are compared to determine the validity of each of the modeling approaches.

Experimental and numerical investigations on PDE performance augmentation by means of an ejector

NASA Astrophysics Data System (ADS)

Canteins, G.; Franzetti, F.; Zocłońska, E.; Khasainov, B. A.; Zitoun, R.; Desbordes, D.

2006-06-01

To improve the performance of pulse detonation engines, a 48 cm long cylindrical combustion chamber of 5cm internal diameter (i.d.) is fitted with an ejector of constant section. The role of the ejector is (i) to provide partial confinement of the detonation products escaping from the chamber and (ii) to suck in fresh air and then to increase the mass ejected compared to the ejection of burned gases alone. The combustion chamber is fully filled with a stoichiometric ethylene/oxygen mixture at ambient conditions. Three parameters of the ejector are varied: the i.d. D, the length L, and the position d relative to the thrust wall of the combustion chamber. For various configurations, the specific impulse ( I sp) is determined in single shot experiments. The maximum operating frequency ( f max) and the maximum thrust are then deduced. I sp is measured by means of the ballistic pendulum method, and f max is derived from the pressure signal recorded on the combustion chamber thrust wall. The addition of an ejector increases the specific impulse up to 60% in the best configuration tested, from 164s without ejector to 260s with ejector. The specific impulse can be represented by a single curve using suitable dimensionless parameters. The thrust results for the main ejector studied ( D = 80mm) indicate an optimal ( L, d) configuration that provides a 28% thrust gain. For the same ejector, f max remains constant and equal to the frequency obtained without ejector in a large range of ( L, d) values, before decreasing. Two-dimensional unsteady numerical computations agree reasonably with the experiments, slightly overestimating the experimental values. The results indicate that 80% of the I sp gain comes from the action of the expanding detonation products on the annular end surface of the combustion chamber, governed by the tube wall thickness.

History of Solid Rockets

NASA Technical Reports Server (NTRS)

Green, Becky; Hales, Christy

2017-01-01

Solid rockets were created by accident and their design and uses have evolved over time. Solid rockets are more simple and reliable than liquid rockets, but they have reduced performance capability. All solid rockets have a similar set of failure modes.

Solar-powered compression-enhanced ejector air conditioner

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sokolov, M.; Hershgal, D.

1993-09-01

This article is an extension of an earlier investigation into the possibility of adaptation of the ejector refrigeration cycle to solar air-conditioning. In a previous work the ejector cycle has been proven a viable option only for a limited number of cases. These include systems with combined (heating, cooling, and hot water supply) loads where means for obtaining low condensing temperature are available. The purpose of this work is to extend the applicability of such systems by enhancing their efficiency and thereby improving their economical attractiveness. This is done by introducing the compression enhanced ejector system in which mechanical (rathermore » than thermal) energy is used to boost the pressure of the secondary stream into the ejector, Such a boost improves the performance of the whole system. Similar to the conventional ejector, the compression-enhanced ejector system utilizes practically the same hardware for solar heating during the winter and for solar cooling during the summer. Thus, it is capable of providing a year-round space air-conditioning. Optimization of the best combination in which the solar and refrigeration systems combine through the vapor generator working temperature is also presented.« less

Effect of Operating Frequency on PDE Driven Ejector Thrust Performance

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh; Landry, K.; Shehadeh, R.; Bouvet, N.; Lee, S.-Y.

2005-01-01

Results of an on-going study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) designed to operate at frequencies up to 50 Hz. The PDE used in these experiments utilizes an equi-molar mixture of oxygen and nitrogen as the oxidizer, and ethylene (C2H4) as the fuel, with the propellant mixture having an equivalence ratio of one. A line of sight laser absorption technique was used to determine the time needed for proper filling of the tube. Thrust measurements were made using an integrated spring damper system coupled with a linear variable displacement transducer. The baseline thrust of the PDE was first measured at each desired frequency and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The ejectors had varying lengths, and two different inlet geometries were tested for each ejector configuration. The parameter space for the study included PDE operation frequency, ejector length, overlap distance and the radius of curvature for the ejector inlets. For the studied experimental matrix, the results showed a maximum thrust augmentation of 106% at an operational frequency of 30 Hz.

Study of an engine flow diverter system for a large scale ejector powered aircraft model

NASA Technical Reports Server (NTRS)

Springer, R. J.; Langley, B.; Plant, T.; Hunter, L.; Brock, O.

1981-01-01

Requirements were established for a conceptual design study to analyze and design an engine flow diverter system and to include accommodations for an ejector system in an existing 3/4 scale fighter model equipped with YJ-79 engines. Model constraints were identified and cost-effective limited modification was proposed to accept the ejectors, ducting and flow diverter valves. Complete system performance was calculated and a versatile computer program capable of analyzing any ejector system was developed.

Some tests on small-scale rectangular throat ejector. [thrust augmentation for V/STOL aircraft

NASA Technical Reports Server (NTRS)

Dean, W. N., Jr.; Franke, M. E.

1979-01-01

A small scale rectangular throat ejector with plane slot nozzles and a fixed throat area was tested to determine the effects of diffuser sidewall length, diffuser area ratio, and sidewall nozzle position on thrust and mass augmentation. The thrust augmentation ratio varied from approximately 0.9 to 1.1. Although the ejector did not have good thrust augmentation performance, the effects of the parameters studied are believed to indicate probable trends in thrust augmenting ejectors.

Design and Testing of Scaled Ejector-Diffusers for Jet Engine Test Facility Applications.

DTIC Science & Technology

1983-09-01

the test cell such that the exhaust will be vented into an augmenting tube which acts as an ejector -diffuser assembly. 11 The kinetic energy of the...OF STANDARDS-1963-A ..’I -Dy , - 77 *4********* Z 7.77- NAVAL POSTGRADUATE SCHOOL Monterey, California W I THESIS DESIGN AND TESTING OF SCALED EJECTOR ...PERIOD COVERED Design and Testing of Scaled Ejector - "flglfeerls Thesis~ Diffusers for Jet Engine Test Facility Spebr18 S. PERFORMING ORG. REPORT

Meso-scale controlled motion for a microfluidic drop ejector.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Galambos, Paul C.; Givler, Richard C.; Pohl, Kenneth Roy

2004-12-01

The objective of this LDRD was to develop a uniquely capable, novel droplet solution based manufacturing system built around a new MEMS drop ejector. The development all the working subsystems required was completed, leaving the integration of these subsystems into a working prototype still left to accomplish. This LDRD report will focus on the three main subsystems: (1) MEMS drop ejector--the MEMS ''sideshooter'' effectively ejected 0.25 pl drops at 10 m/s, (2) packaging--a compact ejector package based on a modified EMDIP (Electro-Microfluidic Dual In-line Package--SAND2002-1941) was fabricated, and (3) a vision/stage system allowing precise ejector package positioning in 3 dimensionsmore » above a target was developed.« less

Fluid-acoustic interactions in a low area ratio supersonic jet ejector

NASA Technical Reports Server (NTRS)

Krothapalli, Anjaneyulu; Ross, Christopher; Yamomoto, K.; Joshi, M. C.

1994-01-01

An experimental investigation carried out to determine aerodynamic and acoustic characteristics of a low area ratio rectangular jet ejector is reported. A supersonic primary jet issuing from a rectangular convergent-divergent nozzle of aspect ratio 4, into a rectangular duct of area ratio 3, was used. Improved performance was found when the ejector screech tone is most intense and appears to match the most unstable Strouhal number of the free rectangular jet. When the primary jet was operating at over and ideally expanded conditions, significant noise reduction was obtained with the ejector as compared to a corresponding free jet. Application of particle image velocimetry to high speed ejector flows was demonstrated through the measurement of instantaneous two dimensional velocity fields.

Nonsteady-Flow Thrust Augmenting Ejectors

NASA Technical Reports Server (NTRS)

Foa, J. V.

1979-01-01

Ejector augmenters in which the transfer of mechanical energy from the primary to the secondary flow takes place through the work of interface pressure forces are investigated. Nonsteady flow processes are analyzed from the standpoint of energy transfer efficiency and a comparison of a rotary jet augmenter to an ejector is presented.

Interface concerns of ejector integration in V/STOL aircraft

NASA Technical Reports Server (NTRS)

Lowry, R. B.

1979-01-01

A number of areas which have in the past contributed to weight, complexity, and thrust losses in the ejector-powered V/STOL vehicle were identified. Most of these interfaces taken singly do not represent a severe compromise to the vehicle; however, the bottom line is that the sum of compromises and the subsequent effects on performance, flight operations and maintenance have rendered the ejector V/STOL aircraft unattractive. In addition to some of the unique ejector/aircraft integration problems, the vehicle by virtue of having a V/STOL capability, is compromised in other areas. To be successful and acceptable, the advantages must outweight the disadvantages and simplicity with minimum penalties must be the rule. It is concluded that more emphasis must be placed on the ejector/aircraft interface for the concept to be successful.

Internal Performance of Several Divergent-Shroud Ejector Nozzles with High Divergence Angles

NASA Technical Reports Server (NTRS)

Trout, Arthur M.; Papell, S. Stephen; Povolny, John H.

1957-01-01

Nine divergent-shroud ejector configurations were investigated to determine the effect of shroud divergence angle on ejector internal performance. Unheated dry air was used for both the primary and secondary flows. The decrease in the design-point thrust coefficient with increasing flow divergence angle (angle measured from primary exit to shroud exit) followed very closely a simple relation involving the cosine of the angle. This indicates that design-point thrust performance for divergent-shroud ejectors can be predicted with reasonable accuracy within the range investigated. The decrease in design-point thrust coefficient due to increasing the flow divergence engle from 120deg to 30deg (half-singles) was approximately 6 percent. Ejector air-handling characteristics and the primary-nozzle flow coefficient were not significantly affected by change in shroud divergence angle.

Thrust Augmentation with Mixer/Ejector Systems

NASA Technical Reports Server (NTRS)

Presz, Walter M., Jr.; Reynolds, Gary; Hunter, Craig

2002-01-01

Older commercial aircraft often exceed FAA (Federal Aviation Administration) sideline noise regulations. The major problem is the jet noise associated with the high exhaust velocities of the low bypass ratio engines on such aircraft. Mixer/ejector exhaust systems can provide a simple means of reducing the jet noise on these aircraft by mixing cool ambient air with the high velocity engine gases before they are exhausted to ambient. This paper presents new information on thrust performance predictions, and thrust augmentation capabilities of mixer/ejectors. Results are presented from the recent development program of the patented Alternating Lobe Mixer Ejector Concept (ALMEC) suppressor system for the Gulfstream GII, GIIB and GIII aircraft. Mixer/ejector performance procedures are presented which include classical control volume analyses, compound compressible flow theory, lobed nozzle loss correlations and state of the art computational fluid dynamic predictions. The mixer/ejector thrust predictions are compared to subscale wind tunnel test model data and actual aircraft flight test measurements. The results demonstrate that a properly designed mixer/ejector noise suppressor can increase effective engine bypass ratio and generate large thrust gains at takeoff conditions with little or no thrust loss at cruise conditions. The cruise performance obtained for such noise suppressor systems is shown to be a strong function of installation effects on the aircraft.

Steam ejector-condenser: stage I of a differential vacuum pumping station

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hanson, C.L.; Alger, T.W.

1981-04-01

A steam ejector-condenser unit was built and tested to produce a 10 Torr (13.3 x 10/sup 2/Pa) vacuum with a 2 cm aperture to the atmosphere. This unit is the first stage of a differential vacuum pumping station that will be used with the Experimental Test Accelerator. The accelerator's electron beam will pass through a series of openings from a high vacuum (5 x 10/sup -6/ Torr) to the atmosphere. The differential system consists of four vacuum pumping units separated by 2 cm-diam apertures. Superheated steam is injected near the final beamline orifice to reduce the quantity of atmospheric airmore » flowing into the steam ejector--condenser unit. The steam ejector in the condenser vessel is open at its center to permit passage of the accelerator beam. Five nozzles mounted in a conical array produce the ejector vacuum of 10 Torr. The ejector exhausts into the condenser and forms a barrier to air flow into the lower pressure region. This feature permits high volume cold trapping and cryopumping of water vapor in the remaining lower-pressure stages. Tests have proven that the steam ejector--condenser is a reliable operating unit and suitable for long-term, steady-state accelerator operation.« less

Performance characteristics of low global warming potential R134a alternative refrigerants in ejector-expansion refrigeration system

NASA Astrophysics Data System (ADS)

Mishra, Shubham; Sarkar, Jahar

2016-12-01

Performance assessment of ejector-expansion vapor compression refrigeration system with eco-friendly R134a alternative refrigerants (R152a, R1234yf, R600a, R600, R290, R161, R32, and propylene) is presented for air-conditioning application. Ejector has been modeled by considering experimental data based correlations of component efficiencies to take care of all irreversibilities. Ejector area ratio has been optimized based on maximum coefficient of performance (COP) for typical air-conditioner operating temperatures. Selected refrigerants have been compared based on area ratio, pressure lift ratio, entrainment ratio, COP, COP improvement and volumetric cooling capacity. Effects of normal boiling point and critical point on the performances have been studied as well. Using ejector as an expansion device, maximum improvement in COP is noted in R1234yf (10.1%), which reduces the COP deviation with R134a (4.5% less in basic cycle and 2.5% less in ejector cycle). Hence, R1234yf seems to be best alternative for ejector expansion system due to its mild flammability and comparable volumetric capacity and cooling COP. refrigerant R161 is superior to R134a in terms of both COP and volumetric cooling capacity, although may be restricted for low capacity application due to its flammability.

Acceleration effects in solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Langhenry, M. T.

1986-01-01

The performance variations due to acceleration loads imposed on spinning solid propellant rocket motors are investigated. The four potentially most significant modes of acceleration-induced phenomena are identified from a study of the literature and modeled. The four modes are a mechanical mode which deals with deformations of the propellant and case: a thermodynamic mode which covers acceleration-induced combustion phenomena; a stress mode which covers the stressed propellant's effect on burn rate; and a gas dynamic mode which deals with changes in gas flow in the chamber and through the nozzle. Simplified models of each mode are developed or taken from the literature and are added to an internal ballistics evaluation computer program. The resulting analysis is the first to include all of the modes. In order to do this an original analysis of the mechanical and stress modes was necessary. However, the analysis shows that the stress mode is not important for the circular perforated grains studied. The other effects are shown to have a significant influence on solid rocket motor performance. The magnitude of the different mode effects are such that one may not be ignored over the others as has been done in the past. The results of the analysis are compared to published rocket motor data. The comparisons indicate an erosive burning effect that is a function of spin rate. A qualitative explanation of the erosive effect is presented.

Internal Mixing Studied for GE/ARL Ejector Nozzle

NASA Technical Reports Server (NTRS)

Zaman, Khairul

2005-01-01

To achieve jet noise reduction goals for the High Speed Civil Transport aircraft, researchers have been investigating the mixer-ejector nozzle concept. For this concept, a primary nozzle with multiple chutes is surrounded by an ejector. The ejector mixes low-momentum ambient air with the hot engine exhaust to reduce the jet velocity and, hence, the jet noise. It is desirable to mix the two streams as fast as possible in order to minimize the length and weight of the ejector. An earlier model of the mixer-ejector nozzle was tested extensively in the Aerodynamic Research Laboratory (ARL) of GE Aircraft Engines at Cincinnati, Ohio. While testing was continuing with later generations of the nozzle, the earlier model was brought to the NASA Lewis Research Center for relatively fundamental measurements. Goals of the Lewis study were to obtain details of the flow field to aid computational fluid dynamics (CFD) efforts and obtain a better understanding of the flow mechanisms, as well as to experiment with mixing enhancement devices, such as tabs. The measurements were made in an open jet facility for cold (unheated) flow without a surrounding coflowing stream.

Investigation of Thrust Augmentation and Acoustic Performance by Ejectors on PDE

NASA Astrophysics Data System (ADS)

Xu, Gui-yang; Weng, Chun-sheng; Li, Ning; Huang, Xiao-long

2016-04-01

Thrust augmentation and acoustic performance of a Pulse Detonation Engine (PDE) with ejector system is experimentally investigated. For these tests the LEjector/DEjector is varied from 1.18 to 4 and the axial placement of the ejector relative to the PDE exhaust is varied from an x/DPDE of -3 to 3. Results from the tests show that the optimum LEjector/DEjector based on thrust augmentation and Overall Sound Pressure Level (OASPL) is found to be 2.61. The divergent ejector performed the best based on thrust augmentation, while the reduction effect for OASPL and Peak Sound Pressure Level (PSPL) at 60° is most prominent for the convergent ejector. The optimum axial position based on thrust augmentation is determined to be x/DPDE = 2, while, x/DPDE = 0 based on OASPL and PSPL.

A combined power and ejector refrigeration cycle for low temperature heat sources

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zheng, B.; Weng, Y.W.

A combined power and ejector refrigeration cycle for low temperature heat sources is under investigation in this paper. The proposed cycle combines the organic Rankine cycle and the ejector refrigeration cycle. The ejector is driven by the exhausts from the turbine to produce power and refrigeration simultaneously. A simulation was carried out to analyze the cycle performance using R245fa as the working fluid. A thermal efficiency of 34.1%, an effective efficiency of 18.7% and an exergy efficiency of 56.8% can be obtained at a generating temperature of 395 K, a condensing temperature of 298 K and an evaporating temperature ofmore » 280 K. Simulation results show that the proposed cycle has a big potential to produce refrigeration and most exergy losses take place in the ejector. (author)« less

A Flight Demonstration of Plasma Rocket Propulsion

NASA Technical Reports Server (NTRS)

Petro, Andrew

1999-01-01

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

An experimental investigation of ejector performance based upon different refrigerants

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chen, S.L.; Yen, J.Y.; Huang, M.C.

1998-12-31

This article experimentally compares the characteristics of different refrigerants as the working fluid in an ejector cooling system. The study covers common refrigerants including R-113, R-114, R-142b, and R-718. The critical choking conditions against the variation of condenser back pressure, the evaporator pressure, and the generator pressure are determined for each refrigerant. The results are compiled into a convenient performance curve and COP chart. These results can serve as an important reference for future design of ejector cooling systems. Finally, this paper presents a comparison of the performances of different refrigerants in an ejector cooling system.

Thrust and pumping characteristics of cylindrical ejectors using afterburning turbojet gas generator

NASA Technical Reports Server (NTRS)

Samanich, N. E.; Huntley, S. C.

1969-01-01

Static tests of cylindrical ejectors having ejector to primary diameter ratios from 1.1 to 1.6 and ejector length to primary nozzle diameter ratios from 0.9 to 2.1 are reported. Power setting of the J85-13 turbojet engine was varied from part power to maximum afterburning. Corrected secondary weight flow ratio was varied from 0.02 to 0.08 over a range of exhaust nozzle pressure ratios from 2.0 to 9.0. Secondary flow temperature rise and pressure drop characteristics through the nacelle secondary flow passage were also obtained.

An overview of current Navy programs to develop thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Green, K. A.

1979-01-01

The primary objective of Navy sponsored research in thrust augmentation is the development of an improved augmenter for V/STOL application. In support of this goal, a data base is being established to provide an accurate prediction capability for use in ejector design. A general technology development of ejectors and associated effects presently is split into the more specific areas of lift and control, since thrust augmenting ejectors may be suitable for both. Research areas examined include advanced diffuser and end wall design; advanced primary nozzles; analytic studies; augmenting reaction controls; and nozzle design.

A Numerical and Experimental Study of Ejector Internal Flow Structure and Geometry Modification for Maximized Performance

NASA Astrophysics Data System (ADS)

Falsafioon, Mehdi; Aidoun, Zine; Poirier, Michel

2017-12-01

A wide range of industrial refrigeration systems are good candidates to benefit from the cooling and refrigeration potential of supersonic ejectors. These are thermally activated and can use waste heat recovery from industrial processes where it is abundantly generated and rejected to the environment. In other circumstances low cost heat from biomass or solar energy may also be used in order to produce a cooling effect. Ejector performance is however typically modest and needs to be maximized in order to take full advantage of the simplicity and low cost of the technology. In the present work, the behavior of ejectors with different nozzle exit positions has been investigated using a prototype as well as a CFD model. The prototype was used in order to measure the performance advantages of refrigerant (R-134a) flowing inside the ejector. For the CFD model, it is assumed that the ejectors are axi-symmetric along x-axis, thus the generated model is in 2D. The preliminary CFD results are validated with experimental data over a wide range of conditions and are in good accordance in terms of entrainment and compression ratios. Next, the flow patterns of four different topologies are studied in order to discuss the optimum geometry in term of ejector entrainment improvement. Finally, The numerical simulations were used to find an optimum value corresponding to maximized entrainment ratio for fixed operating conditions.

Exhaust gas treatment in testing nuclear rocket engines

NASA Astrophysics Data System (ADS)

Zweig, Herbert R.; Fischler, Stanley; Wagner, William R.

1993-01-01

With the exception of the last test series of the Rover program, Nuclear Furnace 1, test-reactor and rocket engine hydrogen gas exhaust generated during the Rover/NERVA program was released directly to the atmosphere, without removal of the associated fission products and other radioactive debris. Current rules for nuclear facilities (DOE Order 5480.6) are far more protective of the general environment; even with the remoteness of the Nevada Test Site, introduction of potentially hazardous quantities of radioactive waste into the atmosphere must be scrupulously avoided. The Rocketdyne treatment concept features a diffuser to provide altitude simulation and pressure recovery, a series of heat exchangers to gradually cool the exhaust gas stream to 100 K, and an activated charcoal bed for adsorption of inert gases. A hydrogen-gas fed ejector provides auxiliary pumping for startup and shutdown of the engine. Supplemental filtration to remove particulates and condensed phases may be added at appropriate locations in the system. The clean hydrogen may be exhausted to the atmosphere and flared, or the gas may be condensed and stored for reuse in testing. The latter approach totally isolates the working gas from the environment.

Ejector/liquid ring pump provides <0. 30 mm Hg vacuum for polymerization vessel

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lockwood, A.; Gaines, A.

1982-03-01

Firestone Fibers and Textiles Company, a division of Firestone Tire and Rubber Company, manufactures tire and industrial yarns of polyester and nylon-6. Nylon-6 molding and extrusion resins are also produced at the plant in Hopewell, Virginia. The process for making polyester requires an extremely low vacuum on the polymerization reactor. A consistent polymerization vessel vacuum of 0.3 mm Hg is needed, but the existing vacuum source, a five-stage steam jet ejector, could only provide a 0.5 mm Hg level. Two options were considered when the company decided to replace the original system with a system designed for 0.15 mm Hgmore » with a non-condensible gas load of 10.8 lb/hr. A new five-stage jet ejector system to meet these requirements would use 1395 lb/hr of 100 psig steam. The other option was a hybrid vacuum source composed of a three-stage steam ejector system and a liquid ring vacuum pump that is more energy efficient than ejectors for low vacuum applications. The hybrid system was selected because the three-stage jet ejector would use only 1240 lb/hr of 100 psig steam. The liquid ring vacuum pump would increase the material and installation cost of the system by about $4000, but the savings in steam consumption would pay back the added cost in less than two years. The jet ejector/liquid ring vacuum pump system has provided both the capacity and the extremely low vacuum needed for the polyester polymerization vessel, after making a small modification. The hybrid vacuum source is reliable, requires only routine maintenance, and will contiue to save substantial amounts of steam each year compared to the five-stage steam jet ejector.« less

Aerodynamic performance investigation of advanced mechanical suppressor and ejector nozzle concepts for jet noise reduction

NASA Technical Reports Server (NTRS)

Wagenknecht, C. D.; Bediako, E. D.

1985-01-01

Advanced Supersonic Transport jet noise may be reduced to Federal Air Regulation limits if recommended refinements to a recently developed ejector shroud exhaust system are successfully carried out. A two-part program consisting of a design study and a subscale model wind tunnel test effort conducted to define an acoustically treated ejector shroud exhaust system for supersonic transport application is described. Coannular, 20-chute, and ejector shroud exhaust systems were evaluated. Program results were used in a mission analysis study to determine aircraft takeoff gross weight to perform a nominal design mission, under Federal Aviation Regulation (1969), Part 36, Stage 3 noise constraints. Mission trade study results confirmed that the ejector shroud was the best of the three exhaust systems studied with a significant takeoff gross weight advantage over the 20-chute suppressor nozzle which was the second best.

Stacked Buoyant Payload Launcher

DTIC Science & Technology

2013-05-14

unit, the signal ejector , or through the escape hatch lockout trunk. Each of these deployment methods has disadvantages. [0005] Torpedo tubes are... ejector tube can accommodate payloads approximately three inches in diameter. Thus, payload size is extremely limited. The escape hatch lockout trunk...signal ejector tube. Additionally, the system 10 can launch multiple payloads during one launch sequence, or can provide multiple launches at

Multiple-cycle Simulation of a Pulse Detonation Engine Ejector

NASA Technical Reports Server (NTRS)

Yungster, S.; Perkins, H. D.

2002-01-01

This paper presents the results of a study involving single and multiple-cycle numerical simulations of various PDE-ejector configurations utilizing hydrogen-oxygen mixtures. The objective was to investigate the thrust, impulse and mass flow rate characteristics of these devices. The results indicate that ejector systems can utilize the energy stored in the strong shock wave exiting the detonation tube to augment the impulse obtained from the detonation tube alone. Impulse augmentation ratios of up to 1.9 were achieved. The axial location of the converging-diverging ejectors relative to the end of the detonation tube were shown to affect the performance of the system.

The wide-range ejector flowmeter: calibrated gas evacuation comprising both high and low gas flows.

PubMed

Waaben, J; Brinkløv, M M; Jørgensen, S

1984-11-01

The wide-range ejector flowmeter is an active scavenging system applying calibrated gas removal directly to the anaesthetic circuit. The evacuation rate can be adjusted on the flowmeter under visual control using the calibration scale ranging from 200 ml X min-1 to 151 X min-1. The accuracy of the calibration was tested on three ejector flowmeters at 12 different presettings. The percentage deviation from presetting varied from + 18 to - 19.4 per cent. The ejector flowmeter enables the provision of consistent and accurately calibrated extraction of waste gases and is applicable within a wide range of fresh gas flows.

Recent development of a jet-diffuser ejector

NASA Technical Reports Server (NTRS)

Alperin, M.; Wu, J. J.

1980-01-01

The paper considers thrust augmenting ejectors in which the processes of mixing and diffusion are partly carried out downstream of the ejector solid surfaces. A jet sheet surrounding the periphery of a widely diverging diffuser prevents separation and forms a gaseous, curved surface to provide effective diffuser ratio and additional length for mixing of primary and induced flows. Three-dimensional potential flow methods achieved a large reduction in the length of the associated solid surface; primary nozzle design further reduced the volume required by the jet-diffuser ejectors, resulting in thrust augmentation in excess of two, and an overall length of about 2 1/2 times the throat width.

Does using an ejector chair affect muscle activation patterns in rheumatoid arthritic patients? A preliminary investigation.

PubMed

Munro, B J; Steele, J R

2000-02-01

The present study examined knee and arm extensor muscle activation patterns displayed by 12 elderly female rheumatoid arthritic patients (mean age = 65.5 +/- 8.6 yr) rising from an instrumented Eser ejector chair under four conditions: high seat (540 mm), low seat (450 mm), with and without ejector assistance. Electromyographic (EMG) signals were sampled (1000 Hz) for vastus lateralis (VL), vastus medialis (VM), rectus femoris (RF) and triceps brachii (TB) using a Noraxon Telemyo System (bandwidth 0-340 Hz). Muscle onset, offset and peak activity relative to loss of seat contact (SS), and integrated EMG, were calculated for each muscle burst before SS. A high seat significantly (p < or = 005) decreased VL and TB intensity but did not change muscle activation patterns compared with rising from a low seat. Ejector assistance significantly increased VM and RF burst duration and RF intensity but had no effect on vastii muscle intensity. It was concluded that concerns pertaining to muscle disuse when rising with ejector assistance were unfounded in the present study. However, further research is required to investigate the effects of habitual use of a mechanical ejector device on muscle activation patterns.

Pratt & Whitney 2D Model in LeRC 9 ft x 15 ft Acoustics

NASA Technical Reports Server (NTRS)

Bridges, James; Marino, Jodilyn

1999-01-01

The theory of mixer-ejectors for noise suppression is illustrated in this cartoon. Since jet noise SPL scales as velocity to the eighth power and diameter squared, increasing the jet diameter while lowering its velocity and keeping thrust constant decreases the noise. However, in supersonic craft, the drag penalty for increasing diameter at supersonic cruise makes this option very expensive. One would like to have a large engine during takeoff which could be shrunk during cruise. The retractable ejector is such an expandable engine. If the mixer flow can be expanded to the size of the ejector exit, the noise generated downstream of the ejector will be much less than the small diameter mixer nozzle alone. Of course, this also requires that the noise created in expanding the flow to fill the ejector be absorbed by a liner in the ejector walls so that none of this noise is heard. Since this mixing of internal hot gas and external cold air must take place in as short a distance as possible, the mixer must be very effective and therefore probably much noisier than a simple nozzle.

A pulse-mode two channel rocket photometer

NASA Astrophysics Data System (ADS)

Petkov, N. P.

Benefits of vertical profile measurements of nighttime emission in the upper atmosphere are discussed. The block diagram of a two-channel rocket photometer with a common pulse operating mode for both channels is described. The requirements and features of the basic units are determined.

A Reduced Ejector Equation,

DTIC Science & Technology

Ejectors , * Thrust augmentation , * Thrust augmentor nozzles, *Mathematical models, Equations, Supersonic characteristics, Inlets, Exits, Aerodynamics, Vertical takeoff aircraft, Short takeoff aircraft, Workshops

Functioning efficiency of intermediate coolers of multistage steam-jet ejectors of steam turbines

NASA Astrophysics Data System (ADS)

Aronson, K. E.; Ryabchikov, A. Yu.; Brodov, Yu. M.; Zhelonkin, N. V.; Murmanskii, I. B.

2017-03-01

Designs of various types of intermediate coolers of multistage ejectors are analyzed and thermal effectiveness and gas-dynamic resistance of coolers are estimated. Data on quantity of steam condensed from steam-air mixture in stage I of an ejector cooler was obtained on the basis of experimental results. It is established that the amount of steam condensed in the cooler constitutes 0.6-0.7 and is almost independent of operating steam pressure (and, consequently, of steam flow) and air amount in steam-air mixture. It is suggested to estimate the amount of condensed steam in a cooler of stage I based on comparison of computed and experimental characteristics of stage II. Computation taking this hypothesis for main types of mass produced multistage ejectors into account shows that 0.60-0.85 of steam amount should be condensed in stage I of the cooler. For ejectors with "pipe-in-pipe" type coolers (EPO-3-200) and helical coolers (EO-30), amount of condensed steam may reach 0.93-0.98. Estimation of gas-dynamic resistance of coolers shows that resistance from steam side in coolers with built-in and remote pipe bundle constitutes 100-300 Pa. Gas-dynamic resistance of "pipein- pipe" and helical type coolers is significantly higher (3-6 times) compared with pipe bundle. However, performance by "dry" (atmospheric) air is higher for ejectors with relatively high gas-dynamic resistance of coolers than those with low resistance at approximately equal operating flow values of ejectors.

Effect of Geometric Parameters on the Performance of Second Throat Annular Steam Ejectors

DTIC Science & Technology

1991-07-01

Cell Pressure versus Rake Average Exit Pitot Pressure . . . . . . . . . . . 42 15. Baseline Wall Pressure Profiles...diffuser exit plane pitot pressure rake . 2.5.2 Alternate Configurations Six alternate ejector diffuser configurations were tested. A summary of...along the walls of the diffusers to help characterize the flow. The ejector diffuser exit pitot pressure was measured with a 6-probe pitot pressure rake

On the rational design of compressible flow ejectors

NASA Technical Reports Server (NTRS)

Ortwerth, P. J.

1979-01-01

A fluid mechanics review of chemical laser ejectors is presented. The characteristics of ejectors with single and multiple driver nozzles are discussed. Methods to compute an optimized performance map in which secondary Mach number and performance are computed versus mass ratio, to compute the flow distortion at each optimized condition, and to determine the thrust area for the design point to match diffuser impedence are examined.

Peptide synthesis on glass substrate using acoustic droplet ejector.

PubMed

Youngki Choe; Shih-Jui Chen; Eun Sok Kim

2014-03-01

This paper describes the synthesis of a 9-mers-long peptide ladder structure of glycine on a modified glass surface using a nanoliter droplet ejector. To synthesize peptide on a glass substrate, SPOT peptide synthesis protocol was followed with a nozzleless acoustic droplet ejector being used to eject about 300 droplets of preactivated amino acid solution to dispense 60 nL of the solution per mer. The coupling efficiency of each mer was measured with FITC fluorescent tag to be 96%, resulting in net 70% efficiency for the whole 9-mer-long peptide of glycine. Usage of a nanoliter droplet ejector for SPOT peptide synthesis increases the density of protein array on a chip.

A Multidisciplinary Approach to Mixer-Ejector Analysis and Design

NASA Technical Reports Server (NTRS)

Hendricks, Eric, S.; Seidel, Jonathan, A.

2012-01-01

The design of an engine for a civil supersonic aircraft presents a difficult multidisciplinary problem to propulsion system engineers. There are numerous competing requirements for the engine, such as to be efficient during cruise while yet quiet enough at takeoff to meet airport noise regulations. The use of mixer-ejector nozzles presents one possible solution to this challenge. However, designing a mixer-ejector which will successfully address both of these concerns is a difficult proposition. Presented in this paper is an integrated multidisciplinary approach to the analysis and design of these systems. A process that uses several low-fidelity tools to evaluate both the performance and acoustics of mixer-ejectors nozzles is described. This process is further expanded to include system-level modeling of engines and aircraft to determine the effects on mission performance and noise near airports. The overall process is developed in the OpenMDAO framework currently being developed by NASA. From the developed process, sample results are given for a notional mixer-ejector design, thereby demonstrating the capabilities of the method.

Effect of Afterbody-Ejector Configurations on the Performance at Transonic Speeds of a Pylon-Supported Nacelle Model having a Hot-Jet Exhaust

NASA Technical Reports Server (NTRS)

Swihart, John M.; Mercer, Charles E.; Norton, Harry T., Jr.

1959-01-01

An investigation of several afterbody-ejector configurations on a pylon-supported nacelle model has been completed in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05. The propulsive performance of two nacelle afterbodies with low boattailing and long ejector spacing was compared with a configuration corresponding to a turbojet-engine installation having a highly boattailed afterbody with a short ejector. The jet exhaust was simulated with a hydrogen peroxide turbojet simulator. The angle of attack was maintained at 0 deg, and the average Reynolds number based on body length was 20 x 10(exp 6). The results of the investigation indicated that the configuration with a conical afterbody with smooth transition to a 15 deg boattail angle had large beneficial jet effects on afterbody pressure-drag coefficient and had the best thrust-minus-drag performance of the afterbody-ejector configurations investigated.

Simulated flight acoustic investigation of treated ejector effectiveness on advanced mechanical suppresors for high velocity jet noise reduction

NASA Technical Reports Server (NTRS)

Brausch, J. F.; Motsinger, R. E.; Hoerst, D. J.

1986-01-01

Ten scale-model nozzles were tested in an anechoic free-jet facility to evaluate the acoustic characteristics of a mechanically suppressed inverted-velocity-profile coannular nozzle with an accoustically treated ejector system. The nozzle system used was developed from aerodynamic flow lines evolved in a previous contract, defined to incorporate the restraints imposed by the aerodynamic performance requirements of an Advanced Supersonic Technology/Variable Cycle Engine system through all its mission phases. Accoustic data of 188 test points were obtained, 87 under static and 101 under simulated flight conditions. The tests investigated variables of hardwall ejector application to a coannular nozzle with 20-chute outer annular suppressor, ejector axial positioning, treatment application to ejector and plug surfaces, and treatment design. Laser velocimeter, shadowgraph photograph, aerodynamic static pressure, and temperature measurement were acquired on select models to yield diagnositc information regarding the flow field and aerodynamic performance characteristics of the nozzles.

A comparison of homogeneous equilibrium and relaxation model for CO2 expansion inside the two-phase ejector

NASA Astrophysics Data System (ADS)

Palacz, M.; Haida, M.; Smolka, J.; Nowak, A. J.; Hafner, A.

2016-09-01

In this study, the comparison of the accuracy of the homogeneous equilibrium model (HEM) and homogeneous relaxation model (HRM) is presented. Both models were applied to simulate the CO2 expansion inside the two-phase ejectors. Moreover, the mentioned models were implemented in the robust and efficient computational tool ejectorPL. That tool guarantees the fully automated computational process and the repeatable computations for the various ejector shapes and operating conditions. The simulated motive nozzle mass flow rates were compared to the experimentally measured mass flow rates. That comparison was made for both, HEM and HRM. The results showed the unsatisfying fidelity of the HEM for the operating regimes far from the carbon dioxide critical point. On the other hand, the HRM accuracy for such conditions was slightly higher. The approach presented in this paper, showed the limitation of applicability of both two-phase models for the expansion phenomena inside the ejectors.

Numerical prediction of 3-D ejector flows

NASA Technical Reports Server (NTRS)

Roberts, D. W.; Paynter, G. C.

1979-01-01

The use of parametric flow analysis, rather than parametric scale testing, to support the design of an ejector system offers a number of potential advantages. The application of available 3-D flow analyses to the design ejectors can be subdivided into several key elements. These are numerics, turbulence modeling, data handling and display, and testing in support of analysis development. Experimental and predicted jet exhaust for the Boeing 727 aircraft are examined.

Conceptual design, evaluation and research identification for Remote Augmented Propulsive Lift Systems (RALS) with ejectors for VTOL aircraft

NASA Technical Reports Server (NTRS)

Willis, W. S.; Konarski, M.; Sutherland, M. V.

1982-01-01

Ejector concepts for use with a remote augmented lift system (RALS) exhaust nozzle were studied. A number of concepts were considered and three were selected as having the greatest promise of providing the desired aircraft and exhaust gas cooling and lift enhancement. A scale model test program is recommended to explore the effects of the more important parameters on ejector performance.

Noise measurements from an ejector suppressor nozzle in the NASA Lewis 9- by 15-foot low speed wind tunnel

NASA Technical Reports Server (NTRS)

Krejsa, Eugene A.; Cooper, Beth A.; Hall, David G.; Khavaran, Abbas

1990-01-01

Acoustic results are presented of a cooperative nozzle test program between NASA and Pratt and Whitney, conducted in the NASA-Lewis 9 x 15 ft Anechoic Wind Tunnel. The nozzle tested was the P and W Hypermix Nozzle concept, a 2-D lobed mixer nozzle followed by a short ejector section made to promote rapid mixing of the induced ejector nozzle flow. Acoustic and aerodynamic measurements were made to determine the amount of ejector pumping, degree of mixing, and noise reduction achieved. A series of tests were run to verify the acoustic quality of this tunnel. The results indicated that the tunnel test section is reasonably anechoic but that background noise can limit the amount of suppression observed from suppressor nozzles. Also, a possible internal noise was observed in the air supply system. The P and W ejector suppressor nozzle demonstrated the potential of this concept to significantly reduce jet noise. Significant reduction in low frequency noise was achieved by increasing the peak jet noise frequency. This was accomplished by breaking the jet into segments with smaller dimensions than those of the baseline nozzle. Variations in ejector parameters had little effect on the noise for the geometries and the range of temperatures and pressure ratios tested.

Ejector-turbine studies and experimental data. Final report, August 1, 1979-October 31, 1982

DOE Office of Scientific and Technical Information (OSTI.GOV)

Minardi, J.E.; Lawson, M.O.; Krolak, R.V.

1982-11-01

An innovative low-power Rankine turbine concept is described which promises competitive efficiencies, low cost, significant reduction in rpm, low maintenance, and long-life operation over similarly rated turbines. The cycle uses a highly efficient two-fluid ejector which greatly lowers the turbine inlet pressure and temperature. The two-fluid ejector cycle is shown by theoretical studies to be capable of transferring energy at efficiencies in excess of 90% from a high-power flux fluid medium to a low-power flux fluid medium. The volume flow of the thermodynamic fluid can be augmented by as much as one-hundred fold. For very low-power turbine applications this couldmore » result in far-more-favorable turbine sizes and rpm. One major application for this type turbine is the heating and cooling with heat pumps. The concept permits engine cycles that cover an extremely broad range of peak temperatures, including those corresponding to stoichiometric combustion of hydrocarbon fuels, waste heat sources, and solar. Actual test data indicated ejector efficiencies as high as 85%. A two-fluid, ejector turbine was designed and tested. The turbine achieved 94% of design power. Additional data indicated that the ejector attached to the turbine operated on the supersonic branch.« less

Liquid sampling system

DOEpatents

Larson, L.L.

1984-09-17

A conduit extends from a reservoir through a sampling station and back to the reservoir in a closed loop. A jet ejector in the conduit establishes suction for withdrawing liquid from the reservoir. The conduit has a self-healing septum therein upstream of the jet ejector for receiving one end of a double-ended cannula, the other end of which is received in a serum bottle for sample collection. Gas is introduced into the conduit at a gas bleed between the sample collection bottle and the reservoir. The jet ejector evacuates gas from the conduit and the bottle and aspirates a column of liquid from the reservoir at a high rate. When the withdrawn liquid reaches the jet ejector the rate of flow therethrough reduces substantially and the gas bleed increases the pressure in the conduit for driving liquid into the sample bottle, the gas bleed forming a column of gas behind the withdrawn liquid column and interrupting the withdrawal of liquid from the reservoir. In the case of hazardous and toxic liquids, the sample bottle and the jet ejector may be isolated from the reservoir and may be further isolated from a control station containing remote manipulation means for the sample bottle and control valves for the jet ejector and gas bleed. 5 figs.

Liquid sampling system

DOEpatents

Larson, Loren L.

1987-01-01

A conduit extends from a reservoir through a sampling station and back to the reservoir in a closed loop. A jet ejector in the conduit establishes suction for withdrawing liquid from the reservoir. The conduit has a self-healing septum therein upstream of the jet ejector for receiving one end of a double-ended cannula, the other end of which is received in a serum bottle for sample collection. Gas is introduced into the conduit at a gas bleed between the sample collection bottle and the reservoir. The jet ejector evacuates gas from the conduit and the bottle and aspirates a column of liquid from the reservoir at a high rate. When the withdrawn liquid reaches the jet ejector the rate of flow therethrough reduces substantially and the gas bleed increases the pressure in the conduit for driving liquid into the sample bottle, the gas bleed forming a column of gas behind the withdrawn liquid column and interrupting the withdrawal of liquid from the reservoir. In the case of hazardous and toxic liquids, the sample bottle and the jet ejector may be isolated from the reservoir and may be further isolated from a control station containing remote manipulation means for the sample bottle and control valves for the jet ejector and gas bleed.

A kinematic and kinetic analysis of the sit-to-stand transfer using an ejector chair: implications for elderly rheumatoid arthritic patients.

PubMed

Munro, B J; Steele, J R; Bashford, G M; Ryan, M; Britten, N

1998-03-01

Twelve elderly female rheumatoid arthritis patients (mean age = 65.5 +/- 8.6 yr) were assessed rising from an instrumented Eser Ejector chair under four conditions: high seat (540 mm), low seat (450 mm), with and without the ejector mechanism operating. Sagittal plane motion, ground reaction forces, and vertical chair arm rest forces were recorded during each trial with the signals synchronised at initial subject head movement. When rising from a high seat, subjects displayed significantly (p < 0.05) greater time to seat off; greater trunk, knee and ankle angles at seat off; increased ankle angular displacement; decreased knee angular displacement; and decreased total net and normalised arm rest forces compared to rising from a low seat. When rising using the ejector mechanism, time to seat off and trunk and knee angle at seat off significantly increased, whereas trunk and knee angular displacement, and total net and normalised arm rest forces significantly decreased compared to rising unassisted. Regardless of seat height or ejector mechanism use, there were no significant differences in the peak, or time to peak horizontal velocity of the subjects' total body centre of mass, or net knee and ankle moments. It was concluded that increased seat height and use of the ejector mechanism facilitated sit-to-stand transfers performed by elderly female rheumatoid arthritic patients. However, using the ejector chair may be preferred by these patients compared to merely raising seat height because it does not necessitate the use of a footstool, a possible obstacle contributing to falls.

Analysis of experimental characteristics of multistage steam-jet electors of steam turbines

NASA Astrophysics Data System (ADS)

Aronson, K. E.; Ryabchikov, A. Yu.; Brodov, Yu. M.; Brezgin, D. V.; Zhelonkin, N. V.; Murmanskii, I. B.

2017-02-01

A series of questions for specification of physical gas dynamics model in flow range of steam-jet unit and ejector computation methodology, as well as functioning peculiarities of intercoolers, was formulated based on analysis of experimental characteristics of multistage team-jet steam turbines. It was established that coefficient defining position of critical cross-section of injected flow depends on characteristics of the "sound tube" zone. Speed of injected flow within this tube may exceed that of sound, and pressure jumps in work-steam decrease at the same time. Characteristics of the "sound tube" define optimal axial sizes of the ejector. According to measurement results, the part of steam condensing in the first-stage coolant constitutes 70-80% of steam amount supplied into coolant and is almost independent of air content in steam. Coolant efficiency depends on steam pressure defined by operation of steam-jet unit of ejector of the next stage after coolant of steam-jet stage, temperature, and condensing water flow. As a rule, steam entering content of steam-air mixture supplied to coolant is overheated with respect to saturation temperature of steam in the mixture. This should be taken into account during coolant computation. Long-term operation causes changes in roughness of walls of the ejector's mixing chamber. The influence of change of wall roughness on ejector characteristic is similar to the influence of reverse pressure of the steam-jet stage. Until some roughness value, injection coefficient of the ejector stage operating in superlimiting regime hardly changed. After reaching critical roughness, the ejector switches to prelimiting operating regime.

Effect of Installation of Mixer/Ejector Nozzles on the Core Flow Exhaust of High-Bypass-Ratio Turbofan Engines

NASA Technical Reports Server (NTRS)

Harrington, Douglas E.

1998-01-01

The aerospace industry is currently investigating the effect of installing mixer/ejector nozzles on the core flow exhaust of high-bypass-ratio turbofan engines. This effort includes both full-scale engine tests at sea level conditions and subscale tests in static test facilities. Subscale model tests are to be conducted prior to full-scale testing. With this approach, model results can be analyzed and compared with analytical predications. Problem areas can then be identified and design changes made and verified in subscale prior to committing to any final design configurations for engine ground tests. One of the subscale model test programs for the integrated mixer/ejector development was a joint test conducted by the NASA Lewis Research Center and Pratt & Whitney Aircraft. This test was conducted to study various mixer/ejector nozzle configurations installed on the core flow exhaust of advanced, high-bypass-ratio turbofan engines for subsonic, commercial applications. The mixer/ejector concept involves the introduction of largescale, low-loss, streamwise vortices that entrain large amounts of secondary air and rapidly mix it with the primary stream. This results in increased ejector pumping relative to conventional ejectors and in more complete mixing within the ejector shroud. The latter improves thrust performance through the efficient energy exchange between the primary and secondary streams. This experimental program was completed in April 1997 in Lewis' CE-22 static test facility. Variables tested included the nozzle area ratio (A9/A8), which ranged from 1.6 to 3.0. This ratio was varied by increasing or decreasing the nozzle throat area, A8. Primary nozzles tested included both lobed mixers and conical primaries. These configurations were tested with and without an outer shroud, and the shroud position was varied by inserting spacers in it. In addition, data were acquired with and without secondary flow.

Combined Numerical/Analytical Perturbation Solutions of the Navier-Stokes Equations for Aerodynamic (Ejector Nozzle) Flows

NASA Technical Reports Server (NTRS)

DeChant, Lawrence J.

1997-01-01

In spite of the rapid advances in both scalar and parallel computational tools, the large number and breadth of variables involved in aerodynamic systems make the use of parabolized or even boundary layer fluid flow models impractical for both preliminary design and inverse design problems. Given this restriction, we have concluded that reduced or approximate models are an important family of tools for design purposes. This study of a combined perturbation/numerical modeling methodology with an application to ejector-mixer nozzles (shown schematically in the following figure) is nearing completion. The work is being funded by a grant from the NASA Lewis Research Center to Texas A&M University. These ejector-mixer nozzle models are designed to be of use to the High Speed Civil Transport Program and may be adopted by both NASA and industry. A computer code incorporating the ejector-mixer models is under development. This code, the Differential Reduced Ejector/Mixer Analysis (DREA), can be run fast enough to be used as a subroutine or to be called by a design optimization routine. Simplified conservation equations--x-momentum, energy, and mass conservation--are used to define the model. Unlike other preliminary design models, DREA requires minimal empirical input and includes vortical mixing and a fully compressible formulation among other features. DREA is being validated by comparing it with results obtained from open literature and proprietary industry data. Preliminary results for a subsonic ejector and a supersonic ejector are shown. In addition, dedicated experiments have been performed at Texas A&M. These experiments use a hydraulic/gas flow analog to provide information about the inviscid mixing interface structure. Final validation and documentation of this work is expected by May of 1997. However, preliminary versions of DREA can be expected in early 1997. In summary, DREA provides a sufficiently detailed and realistic ejector-mixer nozzle model at a computational cost compatible with preliminary design applications.

Impacts of Venturi Turbulent Mixing on the Size Distributions of Sodium Chloride and Dioctyl-Phthalate Aerosols

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cheng, M-D.

2000-08-23

Internal combustion engines are a major source of airborne particulate matter (PM). The size of the engine PM is in the sub-micrometer range. The number of engine particles per unit volume is high, normally in the range of 10{sup 12} to 10{sup 14}. To measure the size distribution of the engine particles dilution of an aerosol sample is required. A diluter utilizing a venturi ejector mixing technique is commercially available and tested. The purpose of this investigation was to determine if turbulence created by the ejector in the mini-dilutor changes the size of particles passing through it. The results ofmore » the NaCl aerosol experiments show no discernible difference in the geometric mean diameter and geometric standard deviation of particles passing through the ejector. Similar results were found for the DOP particles. The ratio of the total number concentrations before and after the ejector indicates that a dilution ratio of approximately 20 applies equally for DOP and NaCl particles. This indicates the dilution capability of the ejector is not affected by the particle composition. The statistical analysis results of the first and second moments of a distribution indicate that the ejector may not change the major parameters (e.g., the geometric mean diameter and geometric standard deviation) characterizing the size distributions of NaCl and DOP particles. However, when the skewness was examined, it indicates that the ejector modifies the particle size distribution significantly. The ejector could change the skewness of the distribution in an unpredictable and inconsistent manner. Furthermore, when the variability of particle counts in individual size ranges as a result of the ejector is examined, one finds that the variability is greater for DOP particles in the size range of 40-150 nm than for NaCl particles in the size range of 30 to 350 nm. The numbers or particle counts in this size region are high enough that the Poisson counting errors are small (<10%) compared with the tail regions. This result shows that the ejector device could have a higher bin-to-bin counting uncertainty for ''soft'' particles such as DOP than for a solid dry particle like NaCl. The results suggest that it may be difficult to precisely characterize the size distribution of particles ejected from the mini-dilution system if the particle is not solid.« less

A modified homogeneous relaxation model for CO2 two-phase flow in vapour ejector

NASA Astrophysics Data System (ADS)

Haida, M.; Palacz, M.; Smolka, J.; Nowak, A. J.; Hafner, A.; Banasiak, K.

2016-09-01

In this study, the homogenous relaxation model (HRM) for CO2 flow in a two-phase ejector was modified in order to increase the accuracy of the numerical simulations The two- phase flow model was implemented on the effective computational tool called ejectorPL for fully automated and systematic computations of various ejector shapes and operating conditions. The modification of the HRM was performed by a change of the relaxation time and the constants included in the relaxation time equation based on the experimental result under the operating conditions typical for the supermarket refrigeration system. The modified HRM was compared to the HEM results, which were performed based on the comparison of motive nozzle and suction nozzle mass flow rates.

A study of air breathing rockets. 3: Supersonic mode combustors

NASA Astrophysics Data System (ADS)

Masuya, G.; Chinzel, N.; Kudo, K.; Murakami, A.; Komuro, T.; Ishii, S.

An experimental study was made on supersonic mode combustors of an air breathing rocket engine. Supersonic streams of room-temperature air and hot fuel-rich rocket exhaust were coaxially mixed and burned in a concially diverging duct of 2 deg half-angle. The effect of air inlet Mach number and excess air ratio was investigated. Axial wall pressure distribution was measured to calculate one dimensional change of Mach number and stagnation temperature. Calculated results showed that supersonic combustion occurred in the duct. At the exit of the duct, gas sampling and Pitot pressure measurement was made, from which radial distributions of various properties were deduced. The distribution of mass fraction of elements from rocket exhaust showed poor mixing performance in the supersonic mode combustors compared with the previously investigated cylindrical subsonic mode combustors. Secondary combustion efficiency correlated well with the centerline mixing parameter, but not with Annushkin's non-dimensional combustor length. No major effect of air inlet Mach number or excess air ratio was seen within the range of conditions under which the experiment was conducted.

Schlieren Imaging of a Single-Ejector, Multi-Tube Pulsed Detonation Engine (Postprint)

DTIC Science & Technology

2009-01-01

studies have shown the potential of an ejector to almost double the thrust of a pulsed detonation engine ( PDE ) tube [1-3]. Axial misalignment of the... Detonation Research Facility in the Air Force Research Laboratory were used for this study. The PDE utilizes automotive valving to feed up to four... detonation tubes. The damped thrust stand was setup to measure PDE thrust alone for baseline tests or total thrust from ejector and PDE . This

Exhaust Nozzles for Supersonic Flight with Turbojet Engines

NASA Technical Reports Server (NTRS)

Shillito, Thomas B.; Hearth, Donald P.; Cortright, Edgar M.

1956-01-01

Good internal performance over a wide range of flight conditions can be obtained with either a plug nozzle or a variable ejector nozzle that can provide a divergent shroud at high pressure ratios. For both the ejector and the plug nozzle, external flow can sometimes cause serious drag losses and, for some plug-nozzle installations, external flow can cause serious internal performance losses. Plug-nozzle cooling and design of the secondary-air-flow systems for ejectors were also considered .

Effects of Multiple Nozzles on Asymmetric Ejector Performance

NASA Technical Reports Server (NTRS)

Lineberry, D.; Landrum, B.

2005-01-01

This paper presents a comparison of a single nozzle and a dual nozzle strut based ejector. The results are focused on the fluid properties in the ejector duct. The research focused on choking mechanisms, mass flow entrainment, and mixing duct pressure distributions. The two ejectors were tests at equivalent primary mass flow rates. This corresponds to chamber pressures ranging from 100 psi to 900 psi in the single nozzle strut and 50 psi to 450 psi in the dual nozzle strut. Secondary flow was drawn from the lab at atmospheric pressure, and was not controlled. The secondary flow was found to choke at a value of 2.3 lb/s for a primary mass flow rate at approximately 2.1 lb/s for both ejectors. This choke was believed to be a mass addition choke rather than a traditional aerodynamic choke. The mixing duct pressure distribution exhibited two distinct trends at "low pressure" trend and at "high pressure" trend. For the low pressure trend, the mixing length for the ejectors remained fixed around 20 inches, regardless of the chamber pressure. For the higher pressure trend, the mixing length was considerably longer and increased with increasing chamber pressure. At high chamber pressures (high mass flow rates), a supersonic core flow was present at the exit of the duct. For these cases, the two streams did not have time to mix by the end of the duct.

Ejector Noise Suppression with Auxiliary Jet Injection

NASA Technical Reports Server (NTRS)

Berman, Charles H.; Andersen, Otto P., Jr.

1997-01-01

An experimental program to reduce aircraft jet turbulence noise investigated the interaction of small auxiliary jets with a larger main jet. Significant reductions in the far field jet noise were obtained over a range of auxiliary jet pressures and flow rates when used in conjunction with an acoustically lined ejector. While the concept is similar to that of conventional ejector suppressors that use mechanical mixing devices, the present approach should improve thrust and lead to lower weight and less complex noise suppression systems since no hardware needs to be located in the main jet flow. A variety of auxiliary jet and ejector configurations and operating conditions were studied. The best conditions tested produced peak to peak noise reductions ranging from 11 to 16 dB, depending on measurement angle, for auxiliary jet mass flows that were 6.6% of the main jet flow with ejectors that were 8 times the main jet diameter in length. Much larger reductions in noise were found at the original peak frequencies of the unsuppressed jet over a range of far field measurement angles.

The effects of the secondary fluid temperature on the energy transfer in an unsteady ejector with a radial-flow diffuser

NASA Astrophysics Data System (ADS)

Ababneh, Amer Khalil; Jawarneh, Ali M.; Tlilan, Hitham M.; Ababneh, Mohammad K.

2009-11-01

Unsteady ejectors are devices whereby energy is exchanged between directly interacting fluids. Unlike steady ejectors, the mechanism responsible for the energy transfer is reversible in nature and thus higher efficiencies are perceivable. A potential application for PEE is for enhancement in output power per weight as in turbochargers. The unsteady ejector when used as a turbocharger the device is expected to perform under wide range of ambient temperatures. Therefore, it is important to investigate the effects of the temperature of the induced ambient air on the energy transfer. The radial-flow ejector, which usually leads to higher-pressure ratios with fewer stages, was selected for the investigation. The flow field is investigated at two Mach numbers 2.5 and 3.0 utilizing rectangular short-length supersonic nozzles for accelerating the primary fluid. Fundamental to the enhancement of these devices performance relies on the management of the flow field in such a way to minimize entropy production. The numerical analyses were conducted utilizing a package of computational fluid dynamics.

System Regulates the Water Contents of Fuel-Cell Streams

NASA Technical Reports Server (NTRS)

Vasquez, Arturo; Lazaroff, Scott

2005-01-01

An assembly of devices provides for both humidification of the reactant gas streams of a fuel cell and removal of the product water (the water generated by operation of the fuel cell). The assembly includes externally-sensing forward-pressure regulators that supply reactant gases (fuel and oxygen) at variable pressures to ejector reactant pumps. The ejector supply pressures depend on the consumption flows. The ejectors develop differential pressures approximately proportional to the consumption flow rates at constant system pressure and with constant flow restriction between the mixer-outlet and suction ports of the ejectors. For removal of product water from the circulating oxygen stream, the assembly includes a water/gas separator that contains hydrophobic and hydrophilic membranes. The water separator imposes an approximately constant flow restriction, regardless of the quality of the two-phase flow that enters it from the fuel cell. The gas leaving the water separator is nearly 100 percent humid. This gas is returned to the inlet of the fuel cell along with a quantity of dry incoming oxygen, via the oxygen ejector, thereby providing some humidification.

Lobed Mixer Optimization for Advanced Ejector Geometries

NASA Technical Reports Server (NTRS)

Waitz, Ian A.

1996-01-01

The overall objectives are: 1) to pursue analytical, computational, and experimental studies that enhance basic understanding of forced mixing phenomena relevant to supersonic jet noise reduction, and 2) to integrate this enhanced understanding (analytical, computational, and empirical) into a design-oriented model of a mixer-ejector noise suppression system. The work is focused on ejector geometries and flow conditions typical of those being investigated in the NASA High Speed Research Program (HSRP). The research will be carried out in collaboration with the NASA HSRP Nozzle Integrated Technology Development (ITD) Team, and will both contribute to, and benefit from, the results of other HSRP research. The noise suppressor system model that is being developed under this grant is distinct from analytical tools developed by industry because it directly links details of lobe geometry to mixer-ejector performance. In addition, the model provides a 'technology road map to define gaps in the current understanding of various phenomena related to mixer-ejector design and to help prioritize research areas. This report describes research completed in the past year, as well as work proposed for the following year.

Thermodynamic Analysis of a Mixed Refrigerant Ejector Refrigeration Cycle Operating with Two Vapor-liquid Separators

NASA Astrophysics Data System (ADS)

Tan, Yingying; Chen, Youming; Wang, Lin

2018-06-01

A mixed refrigerant ejector refrigeration cycle operating with two-stage vapor-liquid separators (MRERC2) is proposed to obtain refrigeration temperature at -40°C. The thermodynamic investigations on performance of MRERC2 using zeotropic mixture refrigerant R23/R134a are performed, and the comparisons of cycle performance between MRERC2 and MRERC1 (MRERC with one-stage vapor-liquid separator) are conducted. The results show that MRERC2 can achieve refrigeration temperature varying between -23.9°C and -42.0°C when ejector pressure ratio ranges from 1.6 to 2.3 at the generation temperature of 57.3-84.9°C. The parametric analysis indicates that increasing condensing temperature decreases coefficient of performance ( COP) of MRERC2, and increasing ejector pressure ratio and mass fraction of the low boiling point component in the mixed refrigerant can improve COP of MRERC2. The MRERC2 shows its potential in utilizing low grade thermal energy as driving power to obtain low refrigeration temperature for the ejector refrigeration cycle.

A Simple Model of Pulsed Ejector Thrust Augmentation

NASA Technical Reports Server (NTRS)

Wilson, Jack; Deloof, Richard L. (Technical Monitor)

2003-01-01

A simple model of thrust augmentation from a pulsed source is described. In the model it is assumed that the flow into the ejector is quasi-steady, and can be calculated using potential flow techniques. The velocity of the flow is related to the speed of the starting vortex ring formed by the jet. The vortex ring properties are obtained from the slug model, knowing the jet diameter, speed and slug length. The model, when combined with experimental results, predicts an optimum ejector radius for thrust augmentation. Data on pulsed ejector performance for comparison with the model was obtained using a shrouded Hartmann-Sprenger tube as the pulsed jet source. A statistical experiment, in which ejector length, diameter, and nose radius were independent parameters, was performed at four different frequencies. These frequencies corresponded to four different slug length to diameter ratios, two below cut-off, and two above. Comparison of the model with the experimental data showed reasonable agreement. Maximum pulsed thrust augmentation is shown to occur for a pulsed source with slug length to diameter ratio equal to the cut-off value.

Transient Ejector Analysis (TEA) code user's guide

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1993-01-01

A FORTRAN computer program for the semi analytic prediction of unsteady thrust augmenting ejector performance has been developed, based on a theoretical analysis for ejectors. That analysis blends classic self-similar turbulent jet descriptions with control-volume mixing region elements. Division of the ejector into an inlet, diffuser, and mixing region allowed flexibility in the modeling of the physics for each region. In particular, the inlet and diffuser analyses are simplified by a quasi-steady-analysis, justified by the assumption that pressure is the forcing function in those regions. Only the mixing region is assumed to be dominated by viscous effects. The present work provides an overview of the code structure, a description of the required input and output data file formats, and the results for a test case. Since there are limitations to the code for applications outside the bounds of the test case, the user should consider TEA as a research code (not as a production code), designed specifically as an implementation of the proposed ejector theory. Program error flags are discussed, and some diagnostic routines are presented.

Ejector subassembly for dual wall air drilling

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kolle, J.J.

1996-09-01

The dry drilling system developed for the Yucca Mountain Site Characterization Project incorporates a surface vacuum system to prevent drilling air and cuttings from contaminating the borehole wall during coring operations. As the drilling depth increases, however there is a potential for borehole contamination because of the limited volume of air which can be removed by the vacuum system. A feasibility analysis has shown that an ejector subassembly mounted in the drill string above the core barrel could significantly enhance the depth capacity of the dry drilling system. The ejector subassembly would use a portion of the air supplied tomore » the core bit to maintain a vacuum on the hole bottom. The results of a design study including performance testing of laboratory scale ejector simulator are presented here.« less

Feasibility study using large ribbon parachutes, retrorockets, and hydrodynamic attenuation to recover liquid rocket boosters for the Space Shuttle

NASA Technical Reports Server (NTRS)

Pepper, William B.; Wailes, William K.

1989-01-01

A new three-phase approach to recovery of the large liquid rocket boosters being studied for the Space Shuttle is proposed. The concept consists of a cluster of larger ribbon parachutes, retrorockets, and spar mode flotation. The two inert liquid rocket boosters weighing 115,000 lb to 183,000 lb descend from high altitude in a side-on coning attitude to 16,000 ft altitude where a cluster of large ribbon parachutes are deployed. The terminal velocity near water landing is 80 ft/sec. Retrorockets are used to decrease the velocity to about 40 ft/sec. The third phase is opening of the front end of the cylindrical rocket case to allow flooding to cushion impact and allow vertical flotation in the spar mode keeping the four expensive liquid rocket engines dry.

The ejector flowmeter as air/oxygen mixing device. An apparatus providing gas mixtures with adjustable oxygen content for high-flow humidification systems.

PubMed

Christensen, K N; Waaben, J; Jørgensen, S

1980-04-01

The ejector flowmeter is constructed for continuous removal of excess gas from anaesthetic circuits. This instrument can be used as an air/oxygen mixing device for high-flow humidification systems in wards where compressed air is not available. Pure oxygen is used as driving gas through the ejector. A nomogram has been constructed to show the relationship between oxygen driving pressure, inlet of air to the flowmeter, FIO2 and total outflow.

Credit WCT. Photographic copy of photograph, view east southeast across ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit WCT. Photographic copy of photograph, view east southeast across Dd station ejectors showing detail of "Hyprox" steam generator. Note that steam generator is placed above Z-stage ejector; an insulated pipe running between the Dd train rails supplies steam to the Y-Stage ejector. Note emergency eyewash stand at extreme right of view. (JPL negative no. 384-3376, 3 December 1962) - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Performance of solar refrigerant ejector refrigerating machine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Al-Khalidy, N.A.H.

1997-12-31

In this work a detailed analysis for the ideal, theoretical, and experimental performance of a solar refrigerant ejector refrigerating machine is presented. A comparison of five refrigerants to select a desirable one for the system is made. The theoretical analysis showed that refrigerant R-113 is more suitable for use in the system. The influence of the boiler, condenser, and evaporator temperatures on system performance is investigated experimentally in a refrigerant ejector refrigerating machine using R-113 as a working refrigerant.

SST Technology Follow-On Program - Phase 2. Noise Suppressor/Nozzle Development. Volume 5. Advanced Suppressor Concepts and Full Scale Tests

DTIC Science & Technology

1975-03-01

Approved for U.S. Government only. This docu- ment is exempted from public availability be- cause of restrictions imposed by the Export Con- trol Act...Transmittal of this document outside the U.S. Government must have prior approval of the Supersonic Transport Office. 20 Security Classif (of this...linings may be determined by comparing nozzle/ejector radiated noise power levels between the unlined (hardwall) ejector case and the lined ejector

Control-Volume Analysis Of Thrust-Augmenting Ejectors

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1990-01-01

New method of analysis of transient flow in thrust-augmenting ejector based on control-volume formulation of governing equations. Considered as potential elements of propulsion subsystems of short-takeoff/vertical-landing airplanes.

Apparatus and method for suppressing sound in a gas turbine engine powerplant

NASA Technical Reports Server (NTRS)

Wynosky, Thomas A. (Inventor); Mischke, Robert J. (Inventor)

1992-01-01

A method and apparatus for suppressing jet noise in a gas turbine engine powerplant 10 is disclosed. Various construction details are developed for providing sound suppression at sea level take-off operative conditions and not providing sound suppression at cruise operative conditions. In one embodiment, the powerplant 10 has a lobed mixer 152 between a primary flowpath 44 and a second flowpath 46, a diffusion region downstream of the lobed mixer region (first mixing region 76), and a deployable ejector/mixer 176 in the diffusion region which forms a second mixing region 78 having a diffusion flowpath 72 downstream of the ejector/mixer and sound absorbing structure 18 bounding the flowpath throughout the diffusion region. The method includes deploying the ejector/mixer 176 at take-off and stowing the ejector/mixer at cruise.

Thermodynamic analysis of a new dual evaporator CO2 transcritical refrigeration cycle

NASA Astrophysics Data System (ADS)

Abdellaoui, Ezzaalouni Yathreb; Kairouani, Lakdar Kairouani

2017-03-01

In this work, a new dual-evaporator CO2 transcritical refrigeration cycle with two ejectors is proposed. In this new system, we proposed to recover the lost energy of condensation coming off the gas cooler and operate the refrigeration cycle ejector free and enhance the system performance and obtain dual-temperature refrigeration simultaneously. The effects of some key parameters on the thermodynamic performance of the modified cycle are theoretically investigated based on energetic and exergetic analysis. The simulation results for the modified cycle indicate more effective system performance improvement than the single ejector in the CO2 vapor compression cycle using ejector as an expander ranging up to 46%. The exergetic analysis for this system is made. The performance characteristics of the proposed cycle show its promise in dual-evaporator refrigeration system.

Use of an approximate similarity principle for the thermal scaling of a full-scale thrust augmenting ejector

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy S.; Perusek, Gail P.; Ibrahim, Mounir B.

1992-01-01

Full temperature ejector model simulations are expensive, and difficult to implement experimentally. If an approximate similarity principle could be established, properly chosen performance parameters should be similar for both hot and cold flow tests if the initial Mach number and total pressures of the flow field are held constant. Existing ejector data is used to explore the utility of one particular similarity principle; the Munk and Prim similarity principle for isentropic flows. Static performance test data for a full-scale thrust augmenting ejector are analyzed for primary flow temperatures up to 1560 R. At different primary temperatures, exit pressure contours are compared for similarity. A nondimensional flow parameter is then used to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments.

Use of an approximate similarity principle for the thermal scaling of a full-scale thrust augmenting ejector

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy; Perusek, Gail P.; Ibrahim, Mounir

1992-01-01

Full temperature ejector model simulations are expensive, and difficult to implement experimentally. If an approximate similarity principle could be established, properly chosen performance parameters should be similar for both hot and cold flow tests if the initial Mach number and total pressures of the flow field are held constant. Existing ejector data is used to explore the utility of one particular similarity principle; the Munk and Prim similarity principle for isentropic flows. Static performance test data for a full-scale thrust augmenting ejector are analyzed for primary flow temperatures up to 1560 R. At different primary temperatures, exit pressure contours are compared for similarity. A nondimensional flow paramenter is then used to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments.

Transition aerodynamics for 20-percent-scale VTOL unmanned aerial vehicle

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Paulson, John W., Jr.

1993-01-01

An investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to establish a transition data base for an unmanned aerial vehicle utilizing a powered-lift ejector system and to evaluate alterations to the ejector system for improved vehicle performance. The model used in this investigation was a 20-percent-scale, blended-body, arrow-wing configuration with integrated twin rectangular ejectors. The test was conducted from hover through transition conditions with variations in angle of attack, angle of sideslip, free-stream dynamic pressure, nozzle pressure ratio, and model ground height. Force and moment data along with extensive surface pressure data were obtained. A laser velocimeter technique for measuring inlet flow velocities was demonstrated at a single flow condition, and also a low order panel method was successfully used to numerically simulate the ejector inlet flow.

CFD study of ejector flow behavior in a blast furnace gas galvanizing plant

NASA Astrophysics Data System (ADS)

Besagni, Giorgio; Mereu, Riccardo; Inzoli, Fabio

2015-02-01

In recent years, there has been a growing interest toward Blast Furnace Gas (BFG) as a low-grade energy source for industrial furnaces. This paper considers the revamping of a galvanic plant furnace converted to BFG from natural gas. In the design of the new system, the ejector on the exhaust line is a critical component. This paper studies the flow behavior of the ejector using a Computational Fluid Dynamics (CFD) analysis. The CFD model is based on a 3D representation of the ejector, using air and exhaust gases as working fluids. This paper is divided in three parts. In the first part, the galvanic plant used as case study is presented and discussed, in the second part the CFD approach is outlined, and in the third part the CFD approach is validated using experimental data and the numerical results are presented and discussed. Different Reynolds-Averaged Navier-Stokes (RANS) turbulence models ( k-ω SST and k-ɛ Realizable) are evaluated in terms of convergence capability and accuracy in predicting the pressure drop along the ejector. Suggestions for future optimization of the system are also provided.

An Experiment on the Near Flow Field of the GE/ARL Mixer Ejector Nozzle

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.

2004-01-01

This report is a documentation of the results on flowfield surveys for the GE/ARL mixer-ejector nozzle carried out in an open jet facility at NASA Glenn Research Center. The results reported are for cold (unheated) flow without any surrounding co-flowing stream. Distributions of streamwise vorticity as well as turbulent stresses, obtained by hot-wire anemometry, are presented for a low subsonic condition. Pitot probe survey results are presented for nozzle pressure ratios up to 3.5. Flowfields both inside and outside of the ejector are considered. Inside the ejector, the mean velocity distribution exhibits a cellular pattern on the cross sectional plane, originating from the flow through the primary and secondary chutes. With increasing downstream distance an interchange of low velocity regions with adjacent high velocity regions takes place due to the action of the streamwise vortices. At the ejector exit, the velocity distribution is nonuniform at low and high pressure ratios but reasonably uniform at intermediate pressure ratios. The effects of two chevron configurations and a tab configuration on the evolution of the downstream jet are also studied. Compared to the baseline case, minor but noticeable effects are observed on the flowfield.

A novel solar bi-ejector refrigeration system and the performance of the added injector with different structures and operation parameters

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wang, Fei; Shen, Shengqiang

2009-12-15

A novel solar bi-ejector refrigeration system was investigated, whose difference compared to the traditional system is that the circulation pump is replaced by a thermal injector. The new system works more stably and needs less maintenance work than the old one, and the whole system can more fully utilize the solar energy. The mathematical models for calculating the performance of the injector and the whole solar refrigeration system were established. The pressure rise performance of injector under different structure and operation parameters and the performance of solar bi-ejector refrigeration system were studied with R123. The results show that the dischargedmore » pressure of injector is affected by structure dimensions of injector and operation conditions. With increasing generation temperature, the entrainment ratio of ejector becomes better while that of injector becomes worse and the overall thermal efficiency of the solar bi-ejector refrigeration system first increases and then decreases with an optimum value of 0.132 at generation temperature of 105 C, condensation temperature of 35 C and evaporation temperature of 10 C. (author)« less

GRC-2010-C-00635

NASA Image and Video Library

2006-02-15

New testing is underway in the Aero-Acoustic Propulsion Laboratory (AAPL) at NASA's Glenn Research Center. The research focuses on a model called the Highly Variable Cycle Exhaust System -- a 0.17 scale model of an exhaust system that will operate at subsonic, transonic and supersonic exhaust speeds in a future supersonic business jet. The model features ejector doors used at different angles. Researchers are investigating the impact of these ejectors on the resulting acoustic radiation. Here, Steven Sedensky, a mechanical engineer with Jacobs Sverdrup, takes measurements of the ejector door positions.

Experimental study of cleaning aircraft GTE fuel injectors using a vortex ejector

NASA Astrophysics Data System (ADS)

Evdokimov, O. A.; Piralishvili, Sh A.; Veretennikov, S. V.; Elkes, A. A.

2017-11-01

The main ways of cleaning the fuel injectors and the circuits of jet and vortex ejectors used for pumping gas, liquid and two-phase media, as well as for evacuation of enclosed spaces are analyzed. The possibility of organizing the process of pumping the liquid out of the fuel injection manifold secondary circuit using a vortex ejector is shown experimentally. The regimes of manifold evacuation at various inlet liquid pressure values are studied. The technology of carbon cleaning fuel injectors using a washing liquid at various working process parameters is tested.

Experimental investigation of an ejector-powered free-jet facility

NASA Technical Reports Server (NTRS)

Long, Mary JO

1992-01-01

NASA Lewis Research Center's (LeRC) newly developed Nozzle Acoustic Test Rig (NATR) is a large free-jet test facility powered by an ejector system. In order to assess the pumping performance of this ejector concept and determine its sensitivity to various design parameters, a 1/5-scale model of the NATR was built and tested prior to the operation of the actual facility. This paper discusses the results of the 1/5-scale model tests and compares them with the findings from the full-scale tests.

Pulsed Ejector Wave Propogation Test Program

NASA Technical Reports Server (NTRS)

Fernandez, Rene; Slater, John W.; Paxson, Daniel E.

2003-01-01

The development of, and initial test data from, a nondetonating Pulse Detonation Engine (PDE) simulator tested in the NASA Glenn 1 x 1 foot Supersonic Wind Tunnel (SWT) is presented in this paper. The concept is a pulsed ejector driven by the simulated exhaust of a PDE. This pro- gram is applicable to a PDE entombed in a ramjet flowpath, i.e., a PDE combined-cycle propulsion system. The ejector primary flow is a pulsed, uiiderexpanded, supersonic nozzle simulating the supersonic waves ema- nating from a PDE, while the ejector secondary flow is the 1 x 1 foot SWT test section operated at subsonic Mach numbers. The objective is not to study the detonation details, but the wave physics including t,he start- ing vortices, the extent of propagation of the wave front, the reflection of the wave from the secondary flowpath walls, and the timing of these events of a pulsed ejector, and correlate these with Computational Fluid Dynamics (CFD) code predictions. Pulsed ejectors have been shown to result in a 3 to 1 improvement in LID (length-to-diameter) and a near 2 to 1 improvement in thrust augmentation over a steady ejector. This program will also explore the extent of upstream interactions between an inlet and large, periodically applied, backpressures to the inlet as would be present due to combustion tube detonations in a PDE. These interactions could result in inlet unstart or buzz for a supersonic mixed compression inlet. The design of the present experiment entailed the use of an 2-t diagram characteristics code to study the nozzle filling and purging timescales as well as a series of CFD analyses conducted using the WIND code. The WIND code is a general purpose CFD code for solution of the Reynolds averaged Navier-Stokes equations and can be applied to both steady state and time-accurate calculations. The first, proof-of-concept, test entry (spring 2001) pressure distributions shown here indicate the simulation concept was successful and therefore the experimental approach is sound.

Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2015-01-01

Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at highpressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NOx emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8%.

Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2015-01-01

Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at high pressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NO(x) emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8.

Vacuum chamber with a supersonic flow aerodynamic window

DOEpatents

Hanson, Clark L.

1982-01-01

A supersonic flow aerodynamic window, whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

Enhancing the efficiency of a three-lobe roots blower by means of by-passing gas to the working cavity through an ejector adaptor

NASA Astrophysics Data System (ADS)

Vizgalov, S. V.; Volkov, M. V.; Chekushkin, G. N.; Khisameev, I. G.

2017-08-01

A positive displacement Roots blower with two- or three-lobe straight-tooth or twisted rotors demonstrates high performance with small specific dimensions and is used to boost internal combustion engines, aerate tanks of treatment facilities, is employed in air and gas transport systems in the food, petrochemical and metallurgical industry. It is common knowledge that several solutions have been implemented in Roots blower designs with straight-tooth with three-lobe or more-lobes rotors. It is more practical to bypass a portion of the compressed gas to the working cavity of the blower through an ejector. The purpose of developing a mathematical model for a blower working in conjunction with an ejector adapter and the further research is to determine the efficiency of this scheme un-der different discharge pressure conditions and different ejector active flow temperatures (the gas cooling effect before the nozzle).

Vacuum chamber with a supersonic-flow aerodynamic window

DOEpatents

Hanson, C.L.

1980-10-14

A supersonic flow aerodynamic window is disclosed whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

Integrated Technology Assessment Center (ITAC) Update

NASA Technical Reports Server (NTRS)

Taylor, J. L.; Neely, M. A.; Curran, F. M.; Christensen, E. R.; Escher, D.; Lovell, N.; Morris, Charles (Technical Monitor)

2002-01-01

The Integrated Technology Assessment Center (ITAC) has developed a flexible systems analysis framework to identify long-term technology needs, quantify payoffs for technology investments, and assess the progress of ASTP-sponsored technology programs in the hypersonics area. For this, ITAC has assembled an experienced team representing a broad sector of the aerospace community and developed a systematic assessment process complete with supporting tools. Concepts for transportation systems are selected based on relevance to the ASTP and integrated concept models (ICM) of these concepts are developed. Key technologies of interest are identified and projections are made of their characteristics with respect to their impacts on key aspects of the specific concepts of interest. Both the models and technology projections are then fed into the ITAC's probabilistic systems analysis framework in ModelCenter. This framework permits rapid sensitivity analysis, single point design assessment, and a full probabilistic assessment of each concept with respect to both embedded and enhancing technologies. Probabilistic outputs are weighed against metrics of interest to ASTP using a multivariate decision making process to provide inputs for technology prioritization within the ASTP. ITAC program is currently finishing the assessment of a two-stage-to-orbit (TSTO), rocket-based combined cycle (RBCC) concept and a TSTO turbine-based combined cycle (TBCC) concept developed by the team with inputs from NASA. A baseline all rocket TSTO concept is also being developed for comparison. Boeing has recently submitted a performance model for their Flexible Aerospace System Solution for Tomorrow (FASST) concept and the ISAT program will provide inputs for a single-stage-to-orbit (SSTO) TBCC based concept in the near-term. Both of these latter concepts will be analyzed within the ITAC framework over the summer. This paper provides a status update of the ITAC program.

46 CFR 56.50-95 - Overboard discharges and shell connections.

Code of Federal Regulations, 2014 CFR

2014-10-01

... securing both the cover and the valve when the chute is not in use. When ash-ejectors or similar expelling... ash ejector discharge shall be not less than Schedule 80. (h) Where deck drains, soil lines, and...

46 CFR 56.50-95 - Overboard discharges and shell connections.

Code of Federal Regulations, 2013 CFR

2013-10-01

... securing both the cover and the valve when the chute is not in use. When ash-ejectors or similar expelling... ash ejector discharge shall be not less than Schedule 80. (h) Where deck drains, soil lines, and...

46 CFR 56.50-95 - Overboard discharges and shell connections.

Code of Federal Regulations, 2012 CFR

2012-10-01

... securing both the cover and the valve when the chute is not in use. When ash-ejectors or similar expelling... ash ejector discharge shall be not less than Schedule 80. (h) Where deck drains, soil lines, and...

Application Focused Schlieren to Nozzle Ejector Flowfields

NASA Technical Reports Server (NTRS)

Mitchell, L. Kerry; Ponton, Michael K.; Seiner, John M.; Manning, James C.; Jansen, Bernard J.; Lagen, Nicholas T.

1999-01-01

The motivation of the testing was to reduce noise generated by eddy Mach wave emission via enhanced mixing in the jet plume. This was to be accomplished through the use of an ejector shroud, which would bring in cooler ambient fluid to mix with the hotter jet flow. In addition, the contour of the mixer, with its chutes and lobes, would accentuate the merging of the outer and inner flows. The objective of the focused schlieren work was to characterize the mixing performance inside of the ejector. Using flow visualization allowed this to be accomplished in a non-intrusive manner.

Credit WCT. Photographic copy of photograph, view of Test Stand ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit WCT. Photographic copy of photograph, view of Test Stand "D" from the south with tower ejector system in operation during a 1972 engine test. Note steam evolving from Z-stage ejectors atop the interstage condenser in the tower. Note also the "Hyprox" steam generator straddling the Dd ejector train to the right. The new Dy horizontal train has not been erected as of this date. In the distance is Test Stand "E." (JPL negative no. 384-9766-AC, 28 November 1972) - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

A control-volume method for analysis of unsteady thrust augmenting ejector flows

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1988-01-01

A method for predicting transient thrust augmenting ejector characteristics is presented. The analysis blends classic self-similar turbulent jet descriptions with a control volume mixing region discretization to solicit transient effects in a new way. Division of the ejector into an inlet, diffuser, and mixing region corresponds with the assumption of viscous-dominated phenomenon in the latter. Inlet and diffuser analyses are simplified by a quasi-steady analysis, justified by the assumptions that pressure is the forcing function in those regions. Details of the theoretical foundation, the solution algorithm, and sample calculations are given.

Low-Cost Flow Visualization for a Supersonic Ejector

NASA Technical Reports Server (NTRS)

Olden, George W.; Lineberry, David M.; Linn, Christopher A. B.; Landrum, Brian D.; Hawk, Clark W.

2005-01-01

Shadowgraph techniques were applied to the cold flow ejector facility at the Propulsion Research Center at the University of Alabama in Huntsville. The setup for the experiments was relatively simple and was accomplished at very little cost. Series of shadowgraph images were taken of both dual nozzle and single nozzle strut based ejectors operating over a range of chamber pressures. The density gradient patterns in the shadowgraphs were compared to pressure data measured along the top and side walls of the mixing duct. The shadowgraph images showed the presence of barrel shocks emanating from the nozzles which at low pressures terminated in Mach disks and at higher pressures extended beyond the barrel shape and reflected off the walls of the duct. Based on pressure data from previous testing, reflected shocks were expected on the walls of the duct. The shadowgraph images confirmed the locations of these reflected shocks on the top wall of the duct. The shadowgraph images also showed the structure change which correlated to a change in pitch of the ejector noise, and corresponded to a change in trend of the duct wall pressure ratio distributions. The images produced from the setup provided insight into the complex flow behavior inside the ejector duct. In addition, the techniques were a valuable tool as an educational device for students.

Rockets for spin recovery

NASA Technical Reports Server (NTRS)

Whipple, R. D.

1980-01-01

The potential effectiveness of rockets as an auxiliary means for an aircraft to effect recovery from spins was investigated. The advances in rocket technology produced by the space effort suggested that currently available systems might obviate many of the problems encountered in earlier rocket systems. A modern fighter configuration known to exhibit a flat spin mode was selected. An analytical study was made of the thrust requirements for a rocket spin recovery system for the subject configuration. These results were then applied to a preliminary systems study of rocket components appropriate to the problem. Subsequent spin tunnel tests were run to evaluate the analytical results.

Viscoelastic propellant effects on Space Shuttle Dynamics

NASA Technical Reports Server (NTRS)

Bugg, F.

1981-01-01

The program of solid propellant research performed in support of the space shuttle dynamics modeling effort is described. Stiffness, damping, and compressibility of the propellant and the effects of many variables on these properties are discussed. The relationship between the propellant and solid rocket booster dynamics during liftoff and boost flight conditions and the effects of booster vibration and propellant stiffness on free free solid rocket booster modes are described. Coupled modes of the shuttle system and the effect of propellant stiffness on the interfaces of the booster and the external tank are described. A finite shell model of the solid rocket booster was developed.

Reusable rocket engine intelligent control system framework design, phase 2

NASA Technical Reports Server (NTRS)

Nemeth, ED; Anderson, Ron; Ols, Joe; Olsasky, Mark

1991-01-01

Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.

12. Sewage Ejector Pumps, view to the southwest. These pumps ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

12. Sewage Ejector Pumps, view to the southwest. These pumps are connected to sewage treatment tanks. - Washington Water Power Clark Fork River Cabinet Gorge Hydroelectric Development, Powerhouse, North Bank of Clark Fork River at Cabinet Gorge, Cabinet, Bonner County, ID

Hydrocarbon fluid, ejector refrigeration system

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kowalski, G.J.; Foster, A.R.

1993-08-31

A refrigeration system is described comprising: a vapor ejector cycle including a working fluid having a property such that entropy of the working fluid when in a saturated vapor state decreases as pressure decreases, the vapor ejector cycle comprising: a condenser located on a common fluid flow path; a diverter located downstream from the condenser for diverting the working fluid into a primary fluid flow path and a secondary fluid flow path parallel to the primary fluid flow path; an evaporator located on the secondary fluid flow path; an expansion device located on the secondary fluid flow path upstream ofmore » the evaporator; a boiler located on the primary fluid flow path parallel to the evaporator for boiling the working fluid, the boiler comprising an axially extending core region having a substantially constant cross sectional area and a porous capillary region surrounding the core region, the core region extending a length sufficient to produce a near sonic velocity saturated vapor; and an ejector having an outlet in fluid communication with the inlet of the condenser and an inlet in fluid communication with the outlet of the evaporator and the outlet of the boiler and in which the flows of the working fluid from the evaporator and the boiler are mixed and the pressure of the working fluid is increased to at least the pressure of the condenser, the ejector inlet, located downstream from the axially extending core region, including a primary nozzle located sufficiently close to the outlet of the boiler to minimize a pressure drop between the boiler and the primary nozzle, the primary nozzle of the ejector including a converging section having an included angle and length preselected to receive the working fluid from the boiler as a near sonic velocity saturated vapor.« less

Results of Aero/Acoustic Tests and Analytical Studies of a Two-Dimensional Eight-Lobe Mixer-Ejector Exhaust Nozzle at Takeoff Conditions

NASA Technical Reports Server (NTRS)

Harrington, Douglas (Technical Monitor); Schweiger, P.; Stern, A.; Gamble, E.; Barber, T.; Chiappetta, L.; LaBarre, R.; Salikuddin, M.; Shin, H.; Majjigi, R.

2005-01-01

Hot flow aero-acoustic tests were conducted with Pratt & Whitney's High-Speed Civil Transport (HSCT) Mixer-Ejector Exhaust Nozzles by General Electric Aircraft Engines (GEAE) in the GEAE Anechoic Freejet Noise Facility (Cell 41) located in Evendale, Ohio. The tests evaluated the impact of various geometric and design parameters on the noise generated by a two-dimensional (2-D) shrouded, 8-lobed, mixer-ejector exhaust nozzle. The shrouded mixer-ejector provides noise suppression by mixing relatively low energy ambient air with the hot, high-speed primary exhaust jet. Additional attenuation was obtained by lining the shroud internal walls with acoustic panels, which absorb acoustic energy generated during the mixing process. Two mixer designs were investigated, the high mixing "vortical" and aligned flow "axial", along with variations in the shroud internal mixing area ratios and shroud length. The shrouds were tested as hardwall or lined with acoustic panels packed with a bulk absorber. A total of 21 model configurations at 1:11.47 scale were tested. The models were tested over a range of primary nozzle pressure ratios and primary exhaust temperatures representative of typical HSCT aero thermodynamic cycles. Static as well as flight simulated data were acquired during testing. A round convergent unshrouded nozzle was tested to provide an acoustic baseline for comparison to the test configurations. Comparisons were made to previous test results obtained with this hardware at NASA Glenn's 9- by 15-foot low-speed wind tunnel (LSWT). Laser velocimetry was used to investigate external as well as ejector internal velocity profiles for comparison to computational predictions. Ejector interior wall static pressure data were also obtained. A significant reduction in exhaust system noise was demonstrated with the 2-D shrouded nozzle designs.

A novel high-temperature ejector-topping power cycle

DOE Office of Scientific and Technical Information (OSTI.GOV)

Freedman, B.Z.; Lior, N.

1994-01-01

A novel, patented topping power cycle is described that takes its energy from a very high-temperature heat source and in which the temperature of the heat sink is still high enough to operate another, conventional power cycle. The top temperatures heat source is used to evaporate a low saturation pressure liquid, which serves as the driving fluid for compressing the secondary fluid in an ejector. Due to the inherently simple construction of ejectors, they are well suited for operation at temperatures higher than those that can be used with gas turbines. The gases exiting from the ejector transfer heat tomore » the lower temperature cycle, and are separated by condensing the primary fluid. The secondary gas is then used to drive a turbine. For a system using sodium as the primary fluid and helium as the secondary fluid, and using a bottoming Rankine steam cycle, the overall thermal efficiency can be at least 11 percent better than that of conventional steam Rankine cycles.« less

Statistical Approaches to Type Determination of the Ejector Marks on Cartridge Cases.

PubMed

Warren, Eric M; Sheets, H David

2018-03-01

While type determination on bullets has been performed for over a century, type determination on cartridge cases is often overlooked. Presented here is an example of type determination of ejector marks on cartridge cases from Glock and Smith & Wesson Sigma series pistols using Naïve Bayes and Random Forest classification methods. The shapes of ejector marks were captured from images of test-fired cartridge cases and subjected to multivariate analysis. Naïve Bayes and Random Forest methods were used to assign the ejector shapes to the correct class of firearm with success rates as high as 98%. This method is easily implemented with equipment already available in crime laboratories and can serve as an investigative lead in the form of a list of firearms that could have fired the evidence. Paired with the FBI's General Rifling Characteristics (GRC) database, this could be an invaluable resource for firearm evidence at crime scenes. © 2017 American Academy of Forensic Sciences.

Jet Noise Reduction Potential from Emerging Variable Cycle Technologies

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Bridges, James; Wernet, Mark

2012-01-01

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation supersonic, commercial aircraft. The concepts were developed by Lockheed Martin (LM), Rolls-Royce Liberty Works (RRLW), and General Electric Global Research (GEGR) as part of an N+2 (next generation forward) aircraft system study initiated by the Supersonics Project in NASA s Fundamental Aeronautics Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at the NASA Glenn Research Center. The exhaust concepts utilized ejectors, inverted velocity profiles, and fluidic shields. One of the ejector concepts was found to produce stagnant flow within the ejector and the other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the model. The concept incorporating an inverted velocity profile and fluid shield produced overall-sound-pressure-level reductions of 6 dB relative to a single stream nozzle at the peak jet noise angle for some nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic performance of the inverted velocity profile model at low nozzle pressure ratios.

Jet Noise Reduction Potential From Emerging Variable Cycle Technologies

NASA Technical Reports Server (NTRS)

2012-01-01

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation supersonic, commercial aircraft. The concepts were developed by Lockheed Martin (LM), Rolls-Royce Liberty Works (RRLW), and General Electric Global Research (GEGR) as part of an N+2 (next generation forward) aircraft system study initiated by the Supersonics Project in NASA s Fundamental Aeronautics Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at the NASA Glenn Research Center. The exhaust concepts utilized ejectors, inverted velocity profiles, and fluidic shields. One of the ejector concepts was found to produce stagnant flow within the ejector and the other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the model. The concept incorporating an inverted velocity profile and fluid shield produced overall-sound-pressure-level reductions of 6 dB relative to a single stream nozzle at the peak jet noise angle for some nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic performance of the inverted velocity profile model at low nozzle pressure ratios.

A three-dimensional turbulent compressible flow model for ejector and fluted mixers

NASA Technical Reports Server (NTRS)

Rushmore, W. L.; Zelazny, S. W.

1978-01-01

A three dimensional finite element computer code was developed to analyze ejector and axisymmetric fluted mixer systems whose flow fields are not significantly influenced by streamwise diffusion effects. A two equation turbulence model was used to make comparisons between theory and data for various flow fields which are components of the ejector system, i.e., (1) turbulent boundary layer in a duct; (2) rectangular nozzle (free jet); (3) axisymmetric nozzle (free jet); (4) hypermixing nozzle (free jet); and (5) plane wall jet. Likewise, comparisons of the code with analytical results and/or other numerical solutions were made for components of the axisymmetric fluted mixer system. These included: (1) developing pipe flow; (2) developing flow in an annular pipe; (3) developing flow in an axisymmetric pipe with conical center body and no fluting and (4) developing fluted pipe flow. Finally, two demonstration cases are presented which show the code's ability to analyze both the ejector and axisymmetric fluted mixers.

Nitrous oxide as a tracer gas in the ASHRAE 110-1995 Standard.

PubMed

Burke, Martin; Wong, Larry; Gonzales, Ben A; Knutson, Gerhard

2014-01-01

ANSI/ASHRAE Standard 110 provides a quantitative method for testing the performance of laboratory fume hoods. Through release of a known quantity (4.0 Lpm) of a tracer gas, and subsequent monitoring of the tracer gas concentration in the "breathing zone" of a mannequin positioned in front of the hood, this method allows for evaluation of laboratory hood performance. Standard 110 specifies sulfur hexafluoride (SF6) as the tracer gas; however, suitable alternatives are allowed. Through three series of performance tests, this analysis serves to investigate the use of nitrous oxide (N2O) as an alternate tracer gas for hood performance testing. Single gas tests were performed according to ASHRAE Standard 110-1995 with each tracer gas individually. These tests showed identical results using an acceptance criterion of AU 0.1 with the sash half open, nominal 18 inches (0.46m) high, and the face velocity at a nominal 60 fpm (0.3 m/s). Most data collected in these single gas tests, for both tracer gases, were below the minimum detection limit, thus two dual gas tests were developed for simultaneous sampling of both tracer gases. Dual gas dual ejector tests were performed with both tracer gases released simultaneously through two ejectors, and the concentration measured with two detectors using a common sampling probe. Dual gas single ejector tests were performed with both tracer gases released though a single ejector, and the concentration measured in the same manner as the dual gas dual ejector tests. The dual gas dual ejector tests showed excellent correlation, with R typically greater than 0.9. Variance was observed in the resulting regression line for each hood, likely due to non-symmetry between the two challenges caused by variables beyond the control of the investigators. Dual gas single ejector tests resulted in exceptional correlation, with R>0.99 typically for the consolidated data, with a slope of 1.0. These data indicate equivalent results for ASHRAE 110 performance testing using either SF6 or N2O, indicating N2O as an applicable alternate tracer gas.

Pressure Regulator With Internal Ejector Circulation Pump, Flow and Pressure Measurement Porting, and Fuel Cell System Integration Options

NASA Technical Reports Server (NTRS)

Vasquez, Arturo

2011-01-01

An advanced reactant pressure regulator with an internal ejector reactant circulation pump has been developed to support NASA's future fuel cell power systems needs. These needs include reliable and safe operation in variable-gravity environments, and for exploration activities with both manned and un manned vehicles. This product was developed for use in Proton Exchange Membrane Fuel Cell (PEMFC) power plant reactant circulation systems, but the design could also be applied to other fuel cell system types, (e.g., solid-oxide or alkaline) or for other gas pressure regulation and circulation needs. The regulator design includes porting for measurement of flow and pressure at key points in the system, and also includes several fuel cell system integration options. NASA has recognized ejectors as a viable alternative to mechanical pumps for use in spacecraft fuel cell power systems. The ejector motive force is provided by a variable, high-pressure supply gas that travels through the ejector s jet nozzle, whereby the pressure energy of the fluid stream is converted to kinetic energy in the gas jet. The ejector can produce circulation-to-consumption-flow ratios that are relatively high (2-3 times), and this phenomenon can potentially (with proper consideration of the remainder of the fuel cell system s design) be used to provide completely for reactant pre-humidification and product water removal in a fuel cell system. Specifically, a custom pressure regulator has been developed that includes: (1) an ejector reactant circulation pump (with interchangeable jet nozzles and mixer sections, gas-tight sliding and static seals in required locations, and internal fluid porting for pressure-sensing at the regulator's control elements) and (2) internal fluid porting to allow for flow rate and system pressure measurements. The fluid porting also allows for inclusion of purge, relief, and vacuum-breaker check valves on the regulator assembly. In addition, this regulator could also be used with NASA's advanced nonflow-through fuel cell power systems by simply incorporating a jet nozzle with an appropriate nozzle diameter.

A User's Guide for the Differential Reduced Ejector/Mixer Analysis "DREA" Program. 1.0

NASA Technical Reports Server (NTRS)

DeChant, Lawrence J.; Nadell, Shari-Beth

1999-01-01

A system of analytical and numerical two-dimensional mixer/ejector nozzle models that require minimal empirical input has been developed and programmed for use in conceptual and preliminary design. This report contains a user's guide describing the operation of the computer code, DREA (Differential Reduced Ejector/mixer Analysis), that contains these mathematical models. This program is currently being adopted by the Propulsion Systems Analysis Office at the NASA Glenn Research Center. A brief summary of the DREA method is provided, followed by detailed descriptions of the program input and output files. Sample cases demonstrating the application of the program are presented.

A modeling technique for STOVL ejector and volume dynamics

NASA Technical Reports Server (NTRS)

Drummond, C. K.; Barankiewicz, W. S.

1990-01-01

New models for thrust augmenting ejector performance prediction and feeder duct dynamic analysis are presented and applied to a proposed Short Take Off and Vertical Landing (STOVL) aircraft configuration. Central to the analysis is the nontraditional treatment of the time-dependent volume integrals in the otherwise conventional control-volume approach. In the case of the thrust augmenting ejector, the analysis required a new relationship for transfer of kinetic energy from the primary flow to the secondary flow. Extraction of the required empirical corrections from current steady-state experimental data is discussed; a possible approach for modeling insight through Computational Fluid Dynamics (CFD) is presented.

Static performance tests of a flight-type STOVL ejector

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy S.

1991-01-01

The design and development of thrust augmenting STOVL ejectors has typically been based on experimental iteration (i.e., trial and error). Static performance tests of a full scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators and yarn tufts) were used to view the inlet air flow, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate seasonal aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature.

Air ejector augmented compressed air energy storage system

DOEpatents

Ahrens, F.W.; Kartsounes, G.T.

Energy is stored in slack demand periods by charging a plurality of underground reservoirs with air to the same peak storage pressure, during peak demand periods throttling the air from one storage reservoir into a gas turbine system at a constant inlet pressure until the air presure in the reservoir falls to said constant inlet pressure, thereupon permitting air in a second reservoir to flow into said gas turbine system while drawing air from the first reservoir through a variable geometry air ejector and adjusting said variable geometry air ejector, said air flow being essentially at the constant inlet pressure of the gas turbine system.

Air ejector augmented compressed air energy storage system

DOEpatents

Ahrens, Frederick W.; Kartsounes, George T.

1980-01-01

Energy is stored in slack demand periods by charging a plurality of underground reservoirs with air to the same peak storage pressure, during peak demand periods throttling the air from one storage reservoir into a gas turbine system at a constant inlet pressure until the air pressure in the reservoir falls to said constant inlet pressure, thereupon permitting air in a second reservoir to flow into said gas turbine system while drawing air from the first reservoir through a variable geometry air ejector and adjusting said variable geometry air ejector, said air flow being essentially at the constant inlet pressure of the gas turbine system.

Transient flow thrust prediction for an ejector propulsion concept

NASA Technical Reports Server (NTRS)

Drummond, Colin K.

1989-01-01

A method for predicting transient thrust augmenting ejector characteristics is introduced. The analysis blends classic self-similar turbulent jet descriptions with a mixing region control volume analysis to predict transient effects in a new way. Details of the theoretical foundation, the solution algorithm, and sample calculations are given.

Jet Engines as High-Capacity Vacuum Pumps

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.

1983-01-01

Large diffuser operations envelope and long run times possible. Jet engine driven ejector/diffuser system combines two turbojet engines and variable-area-ratio ejector in two stages. Applications in such industrial proesses as handling corrosive fumes, evaporation of milk and fruit juices, petroleum distillation, and dehydration of blood plasma and penicillin.

Performance and environmental impact assessment of pulse detonation based engine systems

NASA Astrophysics Data System (ADS)

Glaser, Aaron J.

Experimental research was performed to investigate the feasibility of using pulse detonation based engine systems for practical aerospace applications. In order to carry out this work a new pulse detonation combustion research facility was developed at the University of Cincinnati. This research covered two broad areas of application interest. The first area is pure PDE applications where the detonation tube is used to generate an impulsive thrust directly. The second focus area is on pulse detonation based hybrid propulsion systems. Within each of these areas various studies were performed to quantify engine performance. Comparisons of the performance between detonation and conventional deflagration based engine cycles were made. Fundamental studies investigating detonation physics and flow dynamics were performed in order to gain physical insight into the observed performance trends. Experimental studies were performed on PDE-driven straight and diverging ejectors to determine the system performance. Ejector performance was quantified by thrust measurements made using a damped thrust stand. The effects of PDE operating parameters and ejector geometric parameters on thrust augmentation were investigated. For all cases tested, the maximum thrust augmentation is found to occur at a downstream ejector placement. The optimum ejector geometry was determined to have an overall length of LEJECT/DEJECT =5.61, including an intermediate-straight section length of LSTRT /DEJECT=2, and diverging exhaust section with 4 deg half-angle. A maximum thrust augmentation of 105% was observed while employing the optimized ejector geometry and operating the PDE at a fill-fraction of 0.6 and a frequency of 10 Hz. When operated at a fill-fraction of 1.0 and a frequency of 30 Hz, the thrust augmentation of the optimized PDE-driven ejector system was observed to be 71%. Static pressure was measured along the interior surface of the ejector, including the inlet and exhaust sections. The diverging ejector pressure distribution shows that the diverging section acts as a subsonic diffuser. To provide a better explanation of the observed performance trends, shadowgraph images of the detonation wave and starting vortex interacting with the ejector inlet were obtained. The acoustic signature of a pulse detonation engine was characterized in both the near-field and far-field regimes. Experimental measurements were performed in an anechoic test facility designed for jet noise testing. Both shock strength and speed were mapped as a function of radial distance and direction from the PDE exhaust plane. It was found that the PDE generated pressure field can be reasonably modeled by a theoretical point-source explosion. The effect of several exit nozzle configurations on the PDE acoustic signature was studies. These included various chevron nozzles, a perforated nozzle, and a set of proprietary noise attenuation mufflers. Experimental studies were carried out to investigate the performance of a hybrid propulsion system integrating an axial flow turbine with multiple pulse detonation combustors. The integrated system consisted of a circular array of six pulse detonation combustor (PDC) tubes exhausting through an axial flow turbine. Turbine component performance was quantified by measuring the amount of power generated by the turbine section. Direct comparisons of specific power output and turbine efficiency between a PDC-driven turbine and a turbine driven by steady-flow combustors were made. It was found that the PDC-driven turbine had comparable performance to that of a steady-burner-driven turbine across the operating map of the turbine.

The microspace launcher: first step to the fully air-breathing space launcher

NASA Astrophysics Data System (ADS)

Falempin, F.; Bouchez, M.; Calabro, M.

2009-09-01

A possible application for the high-speed air-breathing propulsion is the fully or partially reusable space launcher. Indeed, by combining the high-speed air-breathing propulsion with a conventional rocket engine (combined cycle or combined propulsion system), it should be possible to improve the average installed specific impulse along the ascent trajectory and then make possible more performing launchers and, hopefully, a fully reusable one. During the last 15 years, a lot of system studies have been performed in France on that subject within the framework of different and consecutive programs. Nevertheless, these studies never clearly demonstrated that a space launcher could take advantage of using a combined propulsion system. During last years, the interest to air-breathing propulsion for space application has been revisited. During this review and taking into account technologies development activities already in progress in Europe, clear priorities have been identified regarding a minimum complementary research and technology program addressing specific needs of space launcher application. It was also clearly identified that there is the need to restart system studies taking advantage of recent progress made regarding knowledge, tools, and technology and focusing on more innovative airframe/propulsion system concepts enabling better trade-off between structural efficiency and propulsion system performance. In that field, a fully axisymmetric configuration has been considered for a microspace launcher (10 kg payload). The vehicle is based on a main stage powered by air-breathing propulsion, combined or not with liquid rocket mode. A "kick stage," powered by a solid rocket engine provides the final acceleration. A preliminary design has been performed for different variants: one using a separated booster and a purely air-breathing main stage, a second one using a booster and a main stage combining air-breathing and rocket mode, a third one without separated booster, the main stage ensuring the initial acceleration in liquid rocket mode and a complementary acceleration phase in rocket mode beyond the air-breathing propulsion system operation. Finally, the liquid rocket engine of this third variant can be replaced by a continuous detonation wave rocket engine. The paper describes the main guidelines for the design of these variants and provides their main characteristics. On this basis, the achievable performance, estimated by trajectory simulation, are detailed.

Credit BG. Looking northwest at the Dd stand complex. To ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. Looking northwest at the Dd stand complex. To the left is the Test Stand "D" tower with steam-driven ejectors and interstage condenser visible along with steam lines. The steam accumulator appears in the left foreground (sphere); steam lines emerging from the top conduct steam to the Dv, Dd, and Dy stand ejectors. The T-shaped vertical pipes atop the accumulator are burst-disk type safety valves. The ejector ends of the Dd and Dy trains are visible to the right. Tracks permitted each train to expand and contract with temperature or equipment changes - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Investigation of the effect of the ejector on the performance of the pulse detonation engine nozzle extension

NASA Astrophysics Data System (ADS)

Korobov, A. E.; Golovastov, S. V.

2015-11-01

Influence of an ejector nozzle extension on gas flow at a pulse detonation engine was investigated numerically and experimentally. Detonation formation was organized in stoichiometric hydrogen-oxygen mixture in cylindrical detonation tube. Cylindrical ejector was constructed and mounted at the open end of the tube. Thrust, air consumption and parameters of the detonation were measured in single and multiple regimes of operation. Axisymmetric model was used in numerical investigation. Equations of Navies-Stokes were solved using a finite-difference scheme Roe of second order of accuracy. Initial conditions were estimated on a base of experimental data. Numerical results were validated with experiments data.

Approximate similarity principle for a full-scale STOVL ejector

NASA Astrophysics Data System (ADS)

Barankiewicz, Wendy S.; Perusek, Gail P.; Ibrahim, Mounir B.

1994-03-01

Full-scale ejector experiments are expensive and difficult to implement at engine exhaust temperatures. For this reason the utility of using similarity principles, in particular the Munk and prim principle for isentropic flow, was explored. Static performance test data for a full-scale thrust augmenting ejector were analyzed for primary flow temperature up to 1560 R. At different primary temperatures, exit pressure contours were compared for similarity. A nondimensional flow parameter is then used to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments. Under the assumption that an appropriate similarity principle can be established, properly chosen performance parameters were found to be similar for both flow and cold flow model tests.

A lifting surface theory for thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Bevilaqua, P. M.

1977-01-01

The circulation theory of airfoil lift has been applied to calculate the performance of thrust augmenting ejectors. The ejector shroud is considered to be 'flying' in the secondary velocity field induced by the entrainment of the primary jet, so that the augmenting thrust is viewed as analogous to the lift on an airfoil. Vortex lattice methods are utilized to compute the thrust augmentation from the force on the flaps. The augmentation is shown to be a function of the length and shape of the flaps, as well as their position and orientation. Predictions of this new theory are compared with the results of classical methods of calculating the augmentation by integration of the stream thrust.

Optical Phenomena Observed upon Some Launches of Russian Rockets

NASA Astrophysics Data System (ADS)

Kozlov, S. I.; Nilolaishvili, S. Sh.; Platov, Yu. V.

2018-01-01

In this paper, unusual optical phenomena observed in our country and abroad upon launches of Russian rockets are discussed and interpreted: they are regarded as the aftereffects of sunlight scattering by gas-dust clouds created by rocket fuel combustion products in different modes of engine operation. The results of instrumental observations of the clouds can be used to study physical processes in the upper atmosphere.

Preliminary engineering report for design of a subscale ejector/diffuser system for high expansion ratio space engine testing

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.; Kurzius, S. C.; Doktor, M. F.

1984-01-01

The design of a subscale jet engine driven ejector/diffuser system is examined. Analytical results and preliminary design drawings and plans are included. Previously developed performance prediction techniques are verified. A safety analysis is performed to determine the mechanism for detonation suppression.

An engine trade study for a supersonic STOVL fighter-attack aircraft, volume 1

NASA Technical Reports Server (NTRS)

Beard, B. B.; Foley, W. H.

1982-01-01

The best main engine for an advanced STOVL aircraft flight demonstrator was studied. The STOVL aircraft uses ejectors powered by engine bypass flow together with vectored core exhaust to achieve vertical thrust capability. Bypass flow and core flow are exhausted through separate nozzles during wingborne flight. Six near term turbofan engines were examined for suitability for this aircraft concept. Fan pressure ratio, thrust split between bypass and core flow, and total thrust level were used to compare engines. One of the six candidate engines was selected for the flight demonstrator configuration. Propulsion related to this aircraft concept was studied. A preliminary candidate for the aircraft reaction control system for hover attitude control was selected. A mathematical model of transfer of bypass thrust from ejectors to aft directed nozzle during the transition to wingborne flight was developed. An equation to predict ejector secondary air flow rate and ram drag is derived. Additional topics discussed include: nozzle area control, ejector to engine inlet reingestion, bypass/core thrust split variation, and gyroscopic behavior during hover.

Reusable Rocket Engine Maintenance Study

NASA Technical Reports Server (NTRS)

Macgregor, C. A.

1982-01-01

Approximately 85,000 liquid rocket engine failure reports, obtained from 30 years of developing and delivering major pump feed engines, were reviewed and screened and reduced to 1771. These were categorized into 16 different failure modes. Failure propagation diagrams were established. The state of the art of engine condition monitoring for in-flight sensors and between flight inspection technology was determined. For the 16 failure modes, the potential measurands and diagnostic requirements were identified, assessed and ranked. Eight areas are identified requiring advanced technology development.

Dynamic characterization of solid rockets

NASA Technical Reports Server (NTRS)

1973-01-01

The structural dynamics of solid rockets in-general was studied. A review is given of the modes of vibration and bending that can exist for a solid propellant rocket, and a NASTRAN computer model is included. Also studied were the dynamic properties of a solid propellant, polybutadiene-acrylic acid-acrylonitrile terpolymer, which may be used in the space shuttle rocket booster. The theory of viscoelastic materials (i.e, Poisson's ratio) was employed in describing the dynamic properties of the propellant. These studies were performed for an eventual booster stage development program for the space shuttle.

Cotton roll isolation versus Vac-Ejector isolation.

PubMed

Wood, A J; Saravia, M E; Farrington, F H

1989-01-01

A visible-light-cured, white pit-and-fissure sealant was applied to 523 teeth in school children using either cotton rolls or a VacEjector for isolation. After a minimum of six months, the patients were recalled and the retention of the sealants was evaluated. No significant difference in sealant retention was found between the two isolation methods.

Credit WCT. Photographic copy of photograph, view west into Dd ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit WCT. Photographic copy of photograph, view west into Dd or Dy ejector, showing steam nozzles which drive the ejector to evacuate the test cell to which it is connected. (JPL negative no. 344-2516-B, 29 August 1977) - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Computer program for calculating the flow field of supersonic ejector nozzles

NASA Technical Reports Server (NTRS)

Anderson, B. H.

1974-01-01

An analytical procedure for computing the performance of supersonic ejector nozzles is presented. This procedure includes real sonic line effects and an interaction analysis for the mixing process between the two streams. The procedure is programmed in FORTRAN 4 and has operated successfully on IBM 7094, IBM 360, CDC 6600, and Univac 1108.

Refrigerant pressurization system with a two-phase condensing ejector

DOEpatents

Bergander, Mark [Madison, CT

2009-07-14

A refrigerant pressurization system including an ejector having a first conduit for flowing a liquid refrigerant therethrough and a nozzle for accelerating a vapor refrigerant therethrough. The first conduit is positioned such that the liquid refrigerant is discharged from the first conduit into the nozzle. The ejector includes a mixing chamber for condensing the vapor refrigerant. The mixing chamber comprises at least a portion of the nozzle and transitions into a second conduit having a substantially constant cross sectional area. The condensation of the vapor refrigerant in the mixing chamber causes the refrigerant mixture in at least a portion of the mixing chamber to be at a pressure greater than that of the refrigerant entering the nozzle and greater than that entering the first conduit.

History of Solid Rockets

NASA Technical Reports Server (NTRS)

Green, Rebecca

2017-01-01

Solid rockets are of interest to the space program because they are commonly used as boosters that provide the additional thrust needed for the space launch vehicle to escape the gravitational pull of the Earth. Larger, more advanced solid rockets allow for space launch vehicles with larger payload capacities, enabling mankind to reach new depths of space. This presentation will discuss, in detail, the history of solid rockets. The history begins with the invention and origin of the solid rocket, and then goes into the early uses and design of the solid rocket. The evolution of solid rockets is depicted by a description of how solid rockets changed and improved and how they were used throughout the 16th, 17th, 18th, and 19th centuries. Modern uses of the solid rocket include the Solid Rocket Boosters (SRBs) on the Space Shuttle and the solid rockets used on current space launch vehicles. The functions and design of the SRB and the advancements in solid rocket technology since the use of the SRB are discussed as well. Common failure modes and design difficulties are discussed as well.

Multiple time scale analysis of pressure oscillations in solid rocket motors

NASA Astrophysics Data System (ADS)

Ahmed, Waqas; Maqsood, Adnan; Riaz, Rizwan

2018-03-01

In this study, acoustic pressure oscillations for single and coupled longitudinal acoustic modes in Solid Rocket Motor (SRM) are investigated using Multiple Time Scales (MTS) method. Two independent time scales are introduced. The oscillations occur on fast time scale whereas the amplitude and phase changes on slow time scale. Hopf bifurcation is employed to investigate the properties of the solution. The supercritical bifurcation phenomenon is observed for linearly unstable system. The amplitude of the oscillations result from equal energy gain and loss rates of longitudinal acoustic modes. The effect of linear instability and frequency of longitudinal modes on amplitude and phase of oscillations are determined for both single and coupled modes. For both cases, the maximum amplitude of oscillations decreases with the frequency of acoustic mode and linear instability of SRM. The comparison of analytical MTS results and numerical simulations demonstrate an excellent agreement.

Dual-fuel, dual-mode rocket engine

NASA Technical Reports Server (NTRS)

Martin, James A. (Inventor)

1989-01-01

The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

Linear quadratic servo control of a reusable rocket engine

NASA Technical Reports Server (NTRS)

Musgrave, Jeffrey L.

1991-01-01

A design method for a servo compensator is developed in the frequency domain using singular values. The method is applied to a reusable rocket engine. An intelligent control system for reusable rocket engines was proposed which includes a diagnostic system, a control system, and an intelligent coordinator which determines engine control strategies based on the identified failure modes. The method provides a means of generating various linear multivariable controllers capable of meeting performance and robustness specifications and accommodating failure modes identified by the diagnostic system. Command following with set point control is necessary for engine operation. A Kalman filter reconstructs the state while loop transfer recovery recovers the required degree of robustness while maintaining satisfactory rejection of sensor noise from the command error. The approach is applied to the design of a controller for a rocket engine satisfying performance constraints in the frequency domain. Simulation results demonstrate the performance of the linear design on a nonlinear engine model over all power levels during mainstage operation.

Space Shuttle Five-Segment Booster (Short Course)

NASA Technical Reports Server (NTRS)

Graves, Stanley R.; Rudolphi, Michael (Technical Monitor)

2002-01-01

NASA is considering upgrading the Space Shuttle by adding a fifth segment (FSB) to the current four-segment solid rocket booster. Course materials cover design and engineering issues related to the Reusable Solid Rocket Motor (RSRM) raised by the addition of a fifth segment to the rocket booster. Topics cover include: four segment vs. five segment booster, abort modes, FSB grain design, erosive burning, enhanced propellant burn rate, FSB erosive burning model development and hardware configuration.

Credit WCT. Photographic copy of photograph, view north across "neutralization ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit WCT. Photographic copy of photograph, view north across "neutralization pond" at Test Stand "D," showing complete Dd station with new Y-Stage and Z-Stage steam-driven ejectors, and "Hyprox" steam generator which powered ejectors. (JPL negative no. 384-3356-B, 20 November 1962) - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Combined cold compressor/ejector helium refrigerator

DOEpatents

Brown, D.P.

1984-06-05

A refrigeration apparatus having an ejector operatively connected with a cold compressor to form a two-stage pumping system. This pumping system is used to lower the pressure, and thereby the temperature of a bath of boiling refrigerant (helium). The apparatus as thus arranged and operated has substantially improved operating efficiency when compared to other processes or arrangements for achieving a similar low pressure.

Combined cold compressor/ejector helium refrigerator

DOEpatents

Brown, Donald P.

1985-01-01

A refrigeration apparatus having an ejector operatively connected with a cold compressor to form a two-stage pumping system. This pumping system is used to lower the pressure, and thereby the temperature of a bath of boiling refrigerant (helium). The apparatus as thus arranged and operated has substantially improved operating efficiency when compared to other processes or arrangements for achieving a similar low pressure.

The ejector-loop fermenter: Description and performance of the apparatus.

PubMed

Moresi, M; Bartolo Gianturco, G; Sebastiani, E

1983-12-01

A novel fermentation unit, the ejector-loop fermenter (ELF), consisting of an outer-loop tower fermenter, a centrifugal pump, a plate-heat exchanger, and a gas-liquid ejector, was designed and constructed. Aeration was achieved by continuously recirculating the fermentation medium through two different nozzle devices instead of using the traditional expensive air compressor. By carrying out a whey fermentation with Kluyveromyces fragilis as the test organism, either in the ELF or in conventional stirred fermenter, it was possible to confirm that the high sheat streses and mixing shock occurring in the ejector nozzle and diffuser sections did not affect microbial growth. Within the range of experimental power consumption per unit volume (-0.1-5 kW/m(3)), the oxygen transfer capability of the ELF per unit power input was found to vary from 1 to 2.5 kg O(2) kW(-1)h(-1). Moreover, it is shown that there is suficient room for improvement in the performance of the ELF unit by care fully designing the aeration device. In fact, at constant volumetric oxygen transfer coefficient, the power consumpotion per unit volume in a 4-mm nozzle was found to be about 40% less than that in a 6-mm nozzle.

A CFD Study of Turbojet and Single-Throat Ramjet Ejector Interaction

NASA Technical Reports Server (NTRS)

Chang, Ing; Hunter, Louis

1996-01-01

Supersonic ejector-diffuse systems have application in driving an advanced airbreathing propulsion system, consisting of turbojet engines acting as the primary and a single throat ramjet acting as the secondary. The turbojet engines are integrated into the single throat ramjet to minimize variable geometry and eliminate redundant propulsion components. The result is a simple, lightweight system that is operable from takeoff to high Mach numbers. At this high Mach number (approximately Mach 3.0), the turbojets are turned off and the high speed ramjet/scramjet take over and drive the vehicle to Mach 6.0. The turbojet-ejector-ramjet system consists of nonafterburning turbojet engines with ducting canted at 20 degrees to supply supersonic flow (downstream of CD nozzle) to the horizontal ramjet duct at a supply total pressure and temperature. Two conditions were modelled by a 2-D full Navier Stokes code at Mach 2.0. The code modelled the Fabri choke as well as the non-Fabri non critical case, using a computational throat to supply the back pressure. The results, which primarily predict the secondary mass flow rate and the mixed conditions at the ejector exit were in reasonable agreement with the 1-D cycle code (TBCC).

An Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wernet, Mark P.; John, Wentworth T.

2004-01-01

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold: first, to add to the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. Second, to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and predicting what thrust augmentation values may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (having no mean component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the various measurements made. These include both thrust and Digital Particle Imaging Velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called Formation Number serves as a useful dimensionless number by which to characterize the jet and predict performance.

Pressure Oscillations and Structural Vibrations in Space Shuttle RSRM and ETM-3 Motors

NASA Technical Reports Server (NTRS)

Mason, D. R.; Morstadt, R. A.; Cannon, S. M.; Gross, E. G.; Nielsen, D. B.

2004-01-01

The complex interactions between internal motor pressure oscillations resulting from vortex shedding, the motor's internal acoustic modes, and the motor's structural vibration modes were assessed for the Space Shuttle four-segment booster Reusable Solid Rocket Motor and for the five-segment engineering test motor ETM-3. Two approaches were applied 1) a predictive procedure based on numerically solving modal representations of a solid rocket motor s acoustic equations of motion and 2) a computational fluid dynamics two-dimensional axi-symmetric large eddy simulation at discrete motor burn times.

Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

Mark, L.

1982-01-01

Conceptual designs and analyses were conducted on two V/STOL supersonic fighter/attack aircraft. These aircraft feature low footprint temperature and pressure thrust augmenting ejectors in the wings for vertical lift, combined with a low wing loading, low wave drag airframe for outstanding cruise and supersonic performance. Aerodynamic, propulsion, performance, and mass properties were determined and are presented for each aircraft. Aerodynamic and Aero/Propulsion characteristics having the most significant effect on the success of the up and away flight mode were identified, and the certainty with which they could be predicted was defined. A wind tunnel model and test program are recommended to resolve the identified uncertainties.

Wet atmospheric generation apparatus

NASA Technical Reports Server (NTRS)

Hamner, Richard M. (Inventor); Allen, Janice K. (Inventor)

1990-01-01

The invention described relates to an apparatus for providing a selectively humidified gas to a camera canister containing cameras and film used in space. A source of pressurized gas (leak test gas or motive gas) is selected by a valve, regulated to a desired pressure by a regulator, and routed through an ejector (venturi device). A regulated source of water vapor in the form of steam from a heated reservoir is coupled to a low pressure region of the ejector which mixes with high velocity gas flow through the ejector. This mixture is sampled by a dew point sensor to obtain dew point thereof (ratio of water vapor to gas) and the apparatus adjusted by varying gas pressure or water vapor to provide a mixture at a connector having selected humidity content.

Cold-flow acoustic evaluation of a small scale, divergent, lobed nozzle for supersonic jet noise suppression

NASA Technical Reports Server (NTRS)

Huff, R. G.; Groesbeck, D. E.

1975-01-01

A supersonic jet noise suppressor was tested with cold flow for acoustic and thrust characteristics at nozzle- to atmospheric-pressure ratios of 1.5 to 4.0. Jet noise suppression and spectral characteristics of the divergent, lobed, suppressor (DLS) nozzle with and without an ejector are presented. Suppression was obtained at nozzle pressure ratios of 2.5 to 4.0. The largest, maximum-lobe, sound pressure level suppression with a hard-wall ejector was 14.6 decibels at a nozzle pressure ratio of 3.5. The thrust loss was 2 percent. In general, low-frequency jet noise was suppressed, leaving higher frequencies essentially unchanged. Without the ejector the nozzle showed a thrust loss of 11 percent together with slightly poorer noise suppression.

Multiple Payload Ejector for Education, Science and Technology Experiments

NASA Technical Reports Server (NTRS)

Lechworth, Gary

2005-01-01

The education research community no longer has a means of being manifested on Space Shuttle flights, and small orbital payload carriers must be flown as secondary payloads on ELV flights, as their launch schedule, secondary payload volume and mass permits. This has resulted in a backlog of small payloads, schedule and cost problems, and an inability for the small payloads community to achieve routine, low-cost access to orbit. This paper will discuss Goddard's Wallops Flight Facility funded effort to leverage its core competencies in small payloads, sounding rockets, balloons and range services to develop a low cost, multiple payload ejector (MPE) carrier for orbital experiments. The goal of the MPE is to provide a low-cost carrier intended primarily for educational flight research experiments. MPE can also be used by academia and industry for science, technology development and Exploration experiments. The MPE carrier will take advantage of the DARPAI NASA partnership to perform flight testing of DARPA s Falcon small, demonstration launch vehicle. The Falcon is similar to MPE fiom the standpoint of focusing on a low-cost, responsive system. Therefore, MPE and Falcon complement each other for the desired long-term goal of providing the small payloads community with a low-cost ride to orbit. The readiness dates of Falcon and MPE are complementary, also. MPE is being developed and readied for flight within 18 months by a small design team. Currently, MPE is preparing for Critical Design Review in fall 2005, payloads are being manifested on the first mission, and the carrier will be ready for flight on the first Falcon demonstration flight in summer, 2006. The MPE and attached experiments can weigh up to 900 lb. to be compatible with Falcon demonstration vehicle lift capabilities fiom Wallops, and will be delivered to the Falcon demonstration orbit - 100 nautical mile circular altitude.

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kubiak, Jonathan M.; Arnett, Lori A.

2016-01-01

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

Dependence of charge transfer phenomena during solid-air two-phase flow on particle disperser

NASA Astrophysics Data System (ADS)

Tanoue, Ken-ichiro; Suedomi, Yuuki; Honda, Hirotaka; Furutani, Satoshi; Nishimura, Tatsuo; Masuda, Hiroaki

2012-12-01

An experimental investigation of the tribo-electrification of particles has been conducted during solid-air two-phase turbulent flow. The current induced in a metal plate by the impact of polymethylmethacrylate (PMMA) particles in a high-speed air flow was measured for two different plate materials. The results indicated that the contact potential difference between the particles and a stainless steel plate was positive, while for a nickel plate it was negative. These results agreed with theoretical contact charge transfer even if not only the particle size but also the kind of metal plate was changed. The specific charge of the PMMA particles during solid-air two-phase flow using an ejector, a stainless steel branch pipe, and a stainless steel straight pipe was measured using a Faraday cage. Although the charge was negative in the ejector, the particles had a positive specific charge at the outlet of the branch pipe, and this positive charge increased in the straight pipe. The charge decay along the flow direction could be reproduced by the charging and relaxation theory. However, the proportional coefficients in the theory changed with the particle size and air velocity. Therefore, an unexpected charge transfer occurred between the ejector and the branch pipe, which could not be explained solely by the contact potential difference. In the ejector, an electrical current in air might have been produced by self-discharge of particles with excess charge between the nickel diffuser in the ejector and the stainless steel nozzle or the stainless steel pipe due to a reversal in the contact potential difference between the PMMA and the stainless steel. The sign of the current depended on the particle size, possibly because the position where the particles impacted depended on their size. When dual coaxial glass pipes were used as a particle disperser, the specific charge of the PMMA particles became more positive along the particle flow direction due to the contact potential difference between the PMMA and the stainless steel. Furthermore, the current in air using the dual coaxial glass pipes was less than that using the ejector.

The analysis of a rocket tomography measurement of the N2+3914A emission and N2 ionization rates in an auroral arc

NASA Technical Reports Server (NTRS)

Mcdade, Ian C.

1991-01-01

Techniques were developed for recovering two-dimensional distributions of auroral volume emission rates from rocket photometer measurements made in a tomographic spin scan mode. These tomographic inversion procedures are based upon an algebraic reconstruction technique (ART) and utilize two different iterative relaxation techniques for solving the problems associated with noise in the observational data. One of the inversion algorithms is based upon a least squares method and the other on a maximum probability approach. The performance of the inversion algorithms, and the limitations of the rocket tomography technique, were critically assessed using various factors such as (1) statistical and non-statistical noise in the observational data, (2) rocket penetration of the auroral form, (3) background sources of emission, (4) smearing due to the photometer field of view, and (5) temporal variations in the auroral form. These tests show that the inversion procedures may be successfully applied to rocket observations made in medium intensity aurora with standard rocket photometer instruments. The inversion procedures have been used to recover two-dimensional distributions of auroral emission rates and ionization rates from an existing set of N2+3914A rocket photometer measurements which were made in a tomographic spin scan mode during the ARIES auroral campaign. The two-dimensional distributions of the 3914A volume emission rates recoverd from the inversion of the rocket data compare very well with the distributions that were inferred from ground-based measurements using triangulation-tomography techniques and the N2 ionization rates derived from the rocket tomography results are in very good agreement with the in situ particle measurements that were made during the flight. Three pre-prints describing the tomographic inversion techniques and the tomographic analysis of the ARIES rocket data are included as appendices.

Ejector Enhanced Pulsejet Based Pressure Gain Combustors: An Old Idea With a New Twist

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Dougherty, Kevin T.

2005-01-01

An experimental investigation of pressure-gain combustion for gas turbine application is described. The test article consists of an off-the-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The combination forms an effective can combustor across which there is a modest total pressure rise rather than the usual loss found in conventional combustors. Although the concept of using a pulsejet to affect semi-constant volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed ejector to efficiently mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure ratio is substantially higher than what has been previously reported in pulsejet-based combustion experiments. Flow non-uniformities in the downstream portion of the device are also shown to be substantially reduced compared to those within the pulsejet itself. The standard deviation of total pressure fluctuations, measured just downstream of the ejector was only 5.0 percent of the mean. This smoothing aspect of the device is critical to turbomachinery applications since turbine performance is, in general, negatively affected by flow non-uniformities and unsteadiness. The experimental rig will be described and details of the performance measurements will be presented. Analyses showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be assessed for several engine types. Issues regarding practical development of such a device are discussed, as are potential emissions reductions resulting from the rich burning nature of the pulsejet and the rapid mixing (quenching) associated with unsteady ejectors.

Optimal coupling and feasibility of a solar-powered year-round ejector air conditioner

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sokolov, M.; Hershgal, D.

1993-06-01

An ejector refrigeration system that uses a conventional refrigerant (R-114) is introduced as a possible mechanism for providing solar-based air-conditioning. Optimal coupling conditions between the collectors' energy output and energy requirements of the cooling system, are investigated. Operation at such optimal conditions assures maximized overall efficiency. Procedures leading to the evaluation of the performance of a real system are disclosed. Design curves for such a system with R-114 as refrigerant are provided. A multi-ejectors arrangement that provides an efficient adjustment for variations of ambient conditions, is described. Year-round air-conditioning is facilitated by rerouting the refrigerant flow through a heating modemore » of the system. Calculations are carried out for illustrative configurations in which relatively low condensing temperature (water reservoirs, cooling towers, or moderate climate) can be maintained.« less

Ejector gas cooling. Phase 1. Final report, 1 April 1987-30 April 1988

DOE Office of Scientific and Technical Information (OSTI.GOV)

MacCracken, C.D.; Silvetti, B.M.; Hrbek, R.

1988-11-01

Closed-circuit ejector cooling systems have never in the past achieved acceptable operating efficiencies in their vapor-compression cycle using standard refrigerants. Despite their long history, relative simplicity, quietness, rugged design, low maintenance and low cost, they could not compete with electric-motor-driven compressors. Phase I is an assessment of two immiscible fluids in an ejector cooling system with different latent heat capacity and molecular weights intended to require less heat in the boiler producing the propellant and taking more heat out in the evaporator cooling fluid. Actual tests corrected to standard conditions and neglecting thermal losses showed 0.5 closed-cycle thermal COP (excludingmore » stack losses), higher than ever previously achieved but below original expectations. Computer programs developed indicate higher COP values are attainable along with competitive first costs.« less

Acoustic and Aerothermal Performance Test of the Axisymmetric Coannular Ejector Nozzle. Volume 2; Acoustic Performance

NASA Technical Reports Server (NTRS)

Herkes, William

2000-01-01

Acoustic and propulsion performance testing of a model-scale Axisymmetric Coannular Ejector nozzle was conducted in the Boeing Low-speed Aeroacoustic Facility. This nozzle is a plug nozzle with an ejector design to provide aspiration of about 20% of the engine flow. A variety of mixing enhancers were designed to promote mixing of the engine and the aspirated flows. These included delta tabs, tone-injection rods, and wheeler ramps. This report addresses the acoustic aspects of the testing. The spectral characteristics of the various configurations of the nozzle are examined on a model-scale basis. This includes indentifying particular noise sources contributing to the spectra and the data are projected to full-scale flyover conditions to evaluate the effectiveness of the nozzle, and of the various mixing enhancers, on reducing the Effective Perceived Noise Levels.

The Role of the Strutjet Engine in New Global and Space Markets

NASA Technical Reports Server (NTRS)

Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

1998-01-01

The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling propulsion concept for single stage to orbit applications. Recent design considerations indicate that this systems also provides benefits supportive of other commercial non-space applications. This paper describes the technical progress of the Strutjet since 1997 together with a rationale why Rocket Based Combined Cycle Engines in general, and the Strutjet in particular, lend themselves uniquely to systems having the ability to expand current space and open new global 'rapid delivery' markets. During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design maturity has been continuously improved and desired features, like simple variable geometry and low drag flowpath resulting in high performance, have been verified. In addition, data is now available which allows the designer, who is challenged to maximize system operability and economic feasibility, to choose between hydrogen or hydrocarbon fuels for a variety of application. The ability exists now to apply this propulsion system to various vehicles with a multitude of missions. In this paper, storable hydrocarbon and gaseous hydrogen Strutjet RBCC test data as accomplished to date and as planned for the future is presented, and the degree of required technology maturity achieved so far is assessed. Two vehicles, using cryogenic propane fuel Strutjet engines, and specifically designed for rapid point-to-point cargo delivery between Pacific rim locations are introduced, discussed, and compared.

Effects of experimentally measured pressure oscillations on the vibration of a solid rocket motor

NASA Technical Reports Server (NTRS)

Schoenster, J. A.; Pierce, H. B.

1972-01-01

Results are presented of firing a Nike rocket against a backstop for the purpose of obtaining pressure fluctuations in the rocket case and determining their relationship to structural vibrations of the case. Special care was required to obtain these pressure fluctuations because of the much higher static pressure generated in the rocket. Very small pressure fluctuations within the rocket case can cause significant vibration levels. A previously observed high frequency was shown to decrease with time before completely disappearing at about 1 second of burning time. The vibration of the case itself is probably related to the longitudinal structural modes at frequencies below 500 Hz and is dependent on local structural conditions at frequencies above this value.

Theoretical and Experimental Analysis of the Constant-Area, Supersonic- Supersonic Ejector

DTIC Science & Technology

1976-10-01

4.3-1 Optimum Chemical Laser System Data, Case No. 1 ....... .... 85 (a) Minn;wum ... ..* ............ 85 ( b ) Minimum P60/P 2...Optimum Chemical Laser System Data, Case No. 2 .... ....... 89 (a) Minimum Wp/Ws .... .. . . . . . . . . . . . . 89 ( b ) Minimum P6 0 /P2...6 Photograph- of the Ejector Model Components .......... ... 48 (a) Front View cf the Secondary Stagnation Chamber * * , 48 ( b ) Rear View of the

This "Is" Rocket Science!

ERIC Educational Resources Information Center

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-01-01

Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

Ejector-Enhanced, Pulsed, Pressure-Gain Combustor

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Dougherty, Kevin T.

2009-01-01

An experimental combination of an off-the-shelf valved pulsejet combustor and an aerodynamically optimized ejector has shown promise as a prototype of improved combustors for gas turbine engines. Despite their name, the constant pressure combustors heretofore used in gas turbine engines exhibit typical pressure losses ranging from 4 to 8 percent of the total pressures delivered by upstream compressors. In contrast, the present ejector-enhanced pulsejet combustor exhibits a pressure rise of about 3.5 percent at overall enthalpy and temperature ratios compatible with those of modern turbomachines. The modest pressure rise translates to a comparable increase in overall engine efficiency and, consequently, a comparable decrease in specific fuel consumption. The ejector-enhanced pulsejet combustor may also offer potential for reducing the emission of harmful exhaust compounds by making it practical to employ a low-loss rich-burn/quench/lean-burn sequence. Like all prior concepts for pressure-gain combustion, the present concept involves an approximation of constant-volume combustion, which is inherently unsteady (in this case, more specifically, cyclic). The consequent unsteadiness in combustor exit flow is generally regarded as detrimental to the performance of downstream turbomachinery. Among other adverse effects, this unsteadiness tends to detract from the thermodynamic benefits of pressure gain. Therefore, it is desirable in any intermittent combustion process to minimize unsteadiness in the exhaust path.

Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wernet, Mark P.; John, Wentworth T.

2007-01-01

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold. The first was to expand the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. The second objective was to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and for predicting the thrust augmentation levels that may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (i.e., it has no mean velocity component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the two main measurements techniques employed. These are thrust and digital particle imaging velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called formation time serves as a useful dimensionless parameter by which to characterize the jet and predict performance.

High-frequency and time resolution rocket observations of structured low- and medium-frequency whistler mode emissions in the auroral ionosphere

NASA Astrophysics Data System (ADS)

LaBelle, J.; McAdams, K. L.; Trimpi, M. L.

High bandwidth electric field waveform measurements on a recent auroral sounding rocket reveal structured whistler mode signals at 400-800 kHz. These are observed intermittently between 300 and 500 km with spectral densities 0-10 dB above the detection threshold of 1.5×10-11V2/m2Hz. The lack of correlation with local particle measurements suggests a remote source. The signals are composed of discrete structures, in one case having bandwidths of about 10 kHz and exhibiting rapid frequency variations of the order of 200 kHz per 100 ms. In one case, emissions near the harmonic of the whistler mode signals are detected simultaneously. Current theories of auroral zone whistler mode emissions have not been applied to explain quantitatively the fine structure of these signals, which resemble auroral kilometric radiation (AKR) rather than auroral hiss.

Vortex Shedding Inside a Baffled Air Duct

NASA Technical Reports Server (NTRS)

Davis, Philip; Kenny, R. Jeremy

2010-01-01

Common in the operation of both segmented and un-segmented large solid rocket motors is the occurrence of vortex shedding within the motor chamber. A portion of the energy within a shed vortex is converted to acoustic energy, potentially driving the longitudinal acoustic modes of the motor in a quasi-discrete fashion. This vortex shedding-acoustic mode excitation event occurs for every Reusable Solid Rocket Motor (RSRM) operation, giving rise to subsequent axial thrust oscillations. In order to better understand this vortex shedding/acoustic mode excitation phenomena, unsteady CFD simulations were run for both a test geometry and the full scale RSRM geometry. This paper covers the results from the subscale geometry runs, which were based on work focusing on the RSRM hydrodynamics. Unsteady CFD simulation parameters, including boundary conditions and post-processing returns, are reviewed. The results were further post-processed to identify active acoustic modes and vortex shedding characteristics. Probable locations for acoustic energy generation, and subsequent acoustic mode excitation, are discussed.

Liquid-propellant rocket engines health-monitoring—a survey

NASA Astrophysics Data System (ADS)

Wu, Jianjun

2005-02-01

This paper is intended to give a summary on the health-monitoring technology, which is one of the key technologies both for improving and enhancing the reliability and safety of current rocket engines and for developing new-generation high reliable reusable rocket engines. The implication of health-monitoring and the fundamental principle obeyed by the fault detection and diagnostics are elucidated. The main aspects of health-monitoring such as system frameworks, failure modes analysis, algorithms of fault detection and diagnosis, control means and advanced sensor techniques are illustrated in some detail. At last, the evolution trend of health-monitoring techniques of liquid-propellant rocket engines is set out.

Failure mode and effects analysis (FMEA) for the Space Shuttle solid rocket motor

NASA Technical Reports Server (NTRS)

Russell, D. L.; Blacklock, K.; Langhenry, M. T.

1988-01-01

The recertification of the Space Shuttle Solid Rocket Booster (SRB) and Solid Rocket Motor (SRM) has included an extensive rewriting of the Failure Mode and Effects Analysis (FMEA) and Critical Items List (CIL). The evolution of the groundrules and methodology used in the analysis is discussed and compared to standard FMEA techniques. Especially highlighted are aspects of the FMEA/CIL which are unique to the analysis of an SRM. The criticality category definitions are presented and the rationale for assigning criticality is presented. The various data required by the CIL and contribution of this data to the retention rationale is also presented. As an example, the FMEA and CIL for the SRM nozzle assembly is discussed in detail. This highlights some of the difficulties associated with the analysis of a system with the unique mission requirements of the Space Shuttle.

Store Separation Simulation of the Penguin Missile from Helicopters

DTIC Science & Technology

2006-05-01

Fin Sections – Parent Aircraft Aerodynamic Modeling • Fuselage • Wing and Pylon – Flight Simulation Features • Eqns. Of Motion • Ejectors , Thrust ...model – Lanyard model – Models for ejectors , thrust , mass, etc… – Helicopter rotor wake model – Penguin wing deployment dynamics – Penguin wing roll...umbilical, wing roll tabs, time dependent thrust and mass properties, and the incorporation of a realistic autopilot. The modeling of the unique

Electro-Microfluidic Packaging

NASA Astrophysics Data System (ADS)

Benavides, G. L.; Galambos, P. C.

2002-06-01

There are many examples of electro-microfluidic products that require cost effective packaging solutions. Industry has responded to a demand for products such as drop ejectors, chemical sensors, and biological sensors. Drop ejectors have consumer applications such as ink jet printing and scientific applications such as patterning self-assembled monolayers or ejecting picoliters of expensive analytes/reagents for chemical analysis. Drop ejectors can be used to perform chemical analysis, combinatorial chemistry, drug manufacture, drug discovery, drug delivery, and DNA sequencing. Chemical and biological micro-sensors can sniff the ambient environment for traces of dangerous materials such as explosives, toxins, or pathogens. Other biological sensors can be used to improve world health by providing timely diagnostics and applying corrective measures to the human body. Electro-microfluidic packaging can easily represent over fifty percent of the product cost and, as with Integrated Circuits (IC), the industry should evolve to standard packaging solutions. Standard packaging schemes will minimize cost and bring products to market sooner.

Gen 2.0 Mixer/Ejector Nozzle Test at LSAF June 1995 to July 1996

NASA Technical Reports Server (NTRS)

Arney, L. D.; Sandquist, D. L.; Forsyth, D. W.; Lidstone, G. L.; Long-Davis, Mary Jo (Technical Monitor)

2005-01-01

Testing of the HSCT Generation 2.0 nozzle model hardware was conducted at the Boeing Low Speed Aeroacoustic Facility, LSAF. Concurrent measurements of noise and thrust were made at critical takeoff design conditions for a variety of mixer/ejector model hardware. Design variables such as suppressor area ratio, mixer area ratio, liner type and thickness, ejector length, lobe penetration, and mixer chute shape were tested. Parallel testing was conducted at G.E.'s Cell 41 acoustic free jet facility to augment the LSAF test. The results from the Gen 2.0 testing are being used to help shape the current nozzle baseline configuration and guide the efforts in the upcoming Generation 2.5 and 3.0 nozzle tests. The Gen 2.0 results have been included in the total airplane system studies conducted at MDC and Boeing to provide updated noise and thrust performance estimates.

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Five-Segment Reusable Solid Rocket Booster Upgrade

NASA Technical Reports Server (NTRS)

Sauvageau, Don

1999-01-01

The Five Segment Reusable Solid Rocket Booster (RSRB) feasibility status is presented in viewgraph form. The Five Segment Booster (FSB) objective is to provide a low cost, low risk approach to increase reliability and safety of the Shuttle system. Topics include: booster upgrade requirements; design summary; reliability issues; booster trajectories; launch site assessment; and enhanced abort modes.

Computational Simulation of Acoustic Modes in Rocket Combustors

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Merkle, C. L.; Sankaran, V.; Ellis, M.

2004-01-01

A combination of computational fluid dynamic analysis and analytical solutions is being used to characterize the dominant modes in liquid rocket engines in conjunction with laboratory experiments. The analytical solutions are based on simplified geometries and flow conditions and are used for careful validation of the numerical formulation. The validated computational model is then extended to realistic geometries and flow conditions to test the effects of various parameters on chamber modes, to guide and interpret companion laboratory experiments in simplified combustors, and to scale the measurements to engine operating conditions. In turn, the experiments are used to validate and improve the model. The present paper gives an overview of the numerical and analytical techniques along with comparisons illustrating the accuracy of the computations as a function of grid resolution. A representative parametric study of the effect of combustor mean flow Mach number and combustor aspect ratio on the chamber modes is then presented for both transverse and longitudinal modes. The results show that higher mean flow Mach numbers drive the modes to lower frequencies. Estimates of transverse wave mechanics in a high aspect ratio combustor are then contrasted with longitudinal modes in a long and narrow combustor to provide understanding of potential experimental simulations.

Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

NASA Technical Reports Server (NTRS)

Morris, C. I.

2003-01-01

Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models.

Effects of damping on mode shapes, volume 1

NASA Technical Reports Server (NTRS)

Gates, R. M.

1977-01-01

Displacement, velocity, and acceleration admittances were calculated for a realistic NASTRAN structural model of space shuttle for three conditions: liftoff, maximum dynamic pressure and end of solid rocket booster burn. The realistic model of the orbiter, external tank, and solid rocket motors included the representation of structural joint transmissibilities by finite stiffness and damping elements. Methods developed to incorporate structural joints and their damping characteristics into a finite element model of the space shuttle, to determine the point damping parameters required to produce realistic damping in the primary modes, and to calculate the effect of distributed damping on structural resonances through the calculation of admittances.

Modal survey of the space shuttle solid rocket motor using multiple input methods

NASA Technical Reports Server (NTRS)

Brillhart, Ralph; Hunt, David L.; Jensen, Brent M.; Mason, Donald R.

1987-01-01

The ability to accurately characterize propellant in a finite element model is a concern of engineers tasked with studying the dynamic response of the Space Shuttle Solid Rocket Motor (SRM). THe uncertainties arising from propellant characterization through specimem testing led to the decision to perform a model survey and model correlation of a single segment of the Shuttle SRM. Multiple input methods were used to excite and define case/propellant modes of both an inert segment and, later, a live propellant segment. These tests were successful at defining highly damped, flexible modes, several pairs of which occured with frequency spacing of less than two percent.

A Preliminary Investigation of Exhaust-Gas Ejectors for Ground Cooling

DTIC Science & Technology

1942-07-01

of teste of ejectore with regard to jet- thrust augmentation . The testa were conducted, for the most; part, with emall- scale models actuated by...their thrust - augmentation , characteristics . D i In view of the lack of experimental data directly applicable to the problem, an invsetigation wae...J e t Propulsion w i t h Special Reference t o Thrust Augmentors. NACA TN No. 442, 1933. 3. Lee, John G.: Progress Report on Augmented J e t

Numerical Evaluation of an Ejector-Enhanced Resonant Pulse Combustor with a Poppet Inlet Valve and a Converging Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2016-01-01

A computational investigation of a pressure-gain combustor system for gas turbine applications is presented. The system consists of a valved pulse combustor and an ejector, housed within a shroud. The study focuses on two enhancements to previous models, related to the valve and ejector components. First, a new poppet inlet valve system is investigated, replacing the previously used reed valve configuration. Secondly, a new computational approach to approximating the effects of choked turbine inlet guide vanes present immediately downstream of the Ejector-Enhanced Resonant Pulse Combustor (EERPC) is investigated. Instead of specifying a back pressure at the EERPC exit boundary (as was done in previous studies) the new model adds a converging-diverging (CD) nozzle at the exit of the EERPC. The throat area of the CD nozzle can be adjusted to obtain the desired back pressure level and total mass flow rate. The results presented indicate that the new poppet valve configuration performs nearly as well as the original reed valve system, and that the addition of the CD nozzle is an effective method to approximate the exit boundary effects of a turbine present downstream of the EERPC. Furthermore, it is shown that the more acoustically reflective boundary imposed by a nozzle as compared to a constant pressure surface does not significantly affect operation or performance.

Probabilistic failure assessment with application to solid rocket motors

NASA Technical Reports Server (NTRS)

Jan, Darrell L.; Davidson, Barry D.; Moore, Nicholas R.

1990-01-01

A quantitative methodology is being developed for assessment of risk of failure of solid rocket motors. This probabilistic methodology employs best available engineering models and available information in a stochastic framework. The framework accounts for incomplete knowledge of governing parameters, intrinsic variability, and failure model specification error. Earlier case studies have been conducted on several failure modes of the Space Shuttle Main Engine. Work in progress on application of this probabilistic approach to large solid rocket boosters such as the Advanced Solid Rocket Motor for the Space Shuttle is described. Failure due to debonding has been selected as the first case study for large solid rocket motors (SRMs) since it accounts for a significant number of historical SRM failures. Impact of incomplete knowledge of governing parameters and failure model specification errors is expected to be important.

Mode Identification of High-Amplitude Pressure Waves in Liquid Rocket Engines

NASA Astrophysics Data System (ADS)

EBRAHIMI, R.; MAZAHERI, K.; GHAFOURIAN, A.

2000-01-01

Identification of existing instability modes from experimental pressure measurements of rocket engines is difficult, specially when steep waves are present. Actual pressure waves are often non-linear and include steep shocks followed by gradual expansions. It is generally believed that interaction of these non-linear waves is difficult to analyze. A method of mode identification is introduced. After presumption of constituent modes, they are superposed by using a standard finite difference scheme for solution of the classical wave equation. Waves are numerically produced at each end of the combustion tube with different wavelengths, amplitudes, and phases with respect to each other. Pressure amplitude histories and phase diagrams along the tube are computed. To determine the validity of the presented method for steep non-linear waves, the Euler equations are numerically solved for non-linear waves, and negligible interactions between these waves are observed. To show the applicability of this method, other's experimental results in which modes were identified are used. Results indicate that this simple method can be used in analyzing complicated pressure signal measurements.

Independent Review of the Failure Modes of F-1 Engine and Propellants System

NASA Technical Reports Server (NTRS)

Ray, Paul

2003-01-01

The F-1 is the powerful engine, that hurdled the Saturn V launch vehicle from the Earth to the moon on July 16,1969. The force that lifted the rocket overcoming the gravitational force during the first stage of the flight was provided by a cluster of five F-1 rocket engines, each of them developing over 1.5 million pounds of thrust (MSFC-MAN-507). The F-1 Rocket engine used RP-1 (Rocket Propellant-1, commercially known as Kerosene), as fuel with lox (liquid Oxygen) as oxidizer. NASA terminated Saturn V activity and has focused on Space Shuttle since 1972. The interest in rocket system has been revived to meet the National Launch System (NLS) program and a directive from the President to return to the Moon and exploration of the space including Mars. The new program Space Launch Initiative (SLI) is directed to drastically reduce the cost of flight for payloads, and adopt a reusable launch vehicle (RLV). To achieve this goal it is essential to have the ability of lifting huge payloads into low earth orbit. Probably requiring powerful boosters as strap-ons to a core vehicle, as was done for the Saturn launch vehicle. The logic in favor of adopting Saturn system, a proven technology, to meet the SLI challenge is very strong. The F-1 engine was the largest and most powerful liquid rocket engine ever built, and had exceptional performance. This study reviews the failure modes of the F-1 engine and propellant system.

Halon Replacement Program for Aviation, Aircraft Engine Nacelle Application Phase II - Operational Comparison of Selected Extinguishants

DTIC Science & Technology

1997-05-01

Control Butterfly Hi-Pressure High Flow Control Butterfly Ejector Primary Clycol Control Valve Scrubber Fan Pressure Control Butterfly 8" Venturi ...the scrubber . 20 ■ SCRUBBER FAN BLOWER INLET VALVE VP-2 VP-3 VP-4 VP-5 VP-6 VP-7 VP-8 VP-9 VP-10 SV-1 SV-2 DESCRIPTION Atmospheric...Blower Bypass Butterfly 24" Venturi Control Butterfly 24" Test Section Exit Butterfly Ejector 10’ Secondary Inlet-Butterfly Hi-pressure Low Flow

Oxygen-iodine ejector laser with a centrifugal bubbling singlet-oxygen generator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zagidullin, M V; Nikolaev, V D; Svistun, M I

2005-10-31

It is shown that if a supersonic oxygen-iodine ejector laser is fed by singlet oxygen from a centrifugal bubbling generator operating at a centrifugal acceleration of {approx}400g, the laser output power achieves a value 1264 W at a chemical efficiency of 24.6% for an alkaline hydrogen peroxide flow rate of 208 cm{sup 3}s{sup -1} and a specific chlorine load of 1.34 mmol s{sup -1} per square centimetre of the bubble layer. (lasers)

Exhaust system for use with a turbine and method of assembling same

DOEpatents

Dalsania, Prakash Bavanjibhai; Sadhu, Antanu

2015-08-18

An exhaust system for use with a steam turbine is provided. An exhaust hood includes an input and an output, the input receiving fluid from the steam turbine. The exhaust hood includes a first side wall that extends between the input and the output. The first side wall includes an aperture. An ejector is coupled to the exhaust hood. The ejector includes inlets and an outlet. At least one of the inlets receives fluid from the exhaust hood via the aperture.

The ejector flowmeter: an evaluation of its accuracy.

PubMed

Waaben, J; Thomsen, A

1978-01-01

The accuracy of five ejector flowmeters was assessed using three different gases and four flow-rates. A soap-bubble flowmeter was used for the calibaration. Significant variations were found between individual flowmeters and between different gas mixtures. No variation was found between the four different flowrates, indicating that the calibration is linear. The mean calibration factor was 84.8% +/- 4.1 (100% O2:87.4 +/- 3.4, 50% N2O/O2: 84.2 +/- 2.8, and 100% N2O: 83.0 +/- 4.6).

Remote mass spectrometric sampling of electrospray- and desorption electrospray-generated ions using an air ejector.

PubMed

Dixon, R Brent; Bereman, Michael S; Muddiman, David C; Hawkridge, Adam M

2007-10-01

A commercial air ejector was coupled to an electrospray ionization linear ion trap mass spectrometer (LTQ) to transport remotely generated ions from both electrospray (ESI) and desorption electrospray ionization (DESI) sources. We demonstrate the remote analysis of a series of analyte ions that range from small molecules and polymers to polypeptides using the AE-LTQ interface. The details of the ESI-AE-LTQ and DESI-AE-LTQ experimental configurations are described and preliminary mass spectrometric data are presented.

Remote Mass Spectrometric Sampling of Electrospray- and Desorption Electrospray-Generated Ions Using an Air Ejector

PubMed Central

Dixon, R. Brent; Bereman, Michael S.; Muddiman, David C.; Hawkridge, Adam M.

2007-01-01

A commercial air ejector was coupled to an electrospray ionization linear ion trap mass spectrometer (LTQ) to transport remotely generated ions from both electrospray (ESI) and desorption electrospray ionization (DESI) sources. We demonstrate the remote analysis of a series of analyte ions that range from small molecules and polymers to polypeptides using the AE-LTQ interface. The details of the ESI-AE-LTQ and DESI-AE-LTQ experimental configurations are described and preliminary mass spectrometric data is presented. PMID:17716909

Diffuser/ejector system for a very high vacuum environment

NASA Technical Reports Server (NTRS)

Riggs, K. E.; Wojciechowski, C. J. (Inventor)

1984-01-01

Turbo jet engines are used to furnish the necessary high temperature, high volume, medium pressure gas to provide a high vacuum test environment at comparatively low cost for space engines at sea level. Moreover, the invention provides a unique way by use of the variable area ratio ejectors with a pair of meshing cones are used. The outer cone is arranged to translate fore and aft, and the inner cone is interchangeable with other cones having varying angles of taper.

Thermal energy recycling fuel cell arrangement

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hanrahan, Paul R.

An example fuel cell arrangement includes a fuel cell stack configured to receive a supply fluid and to provide an exhaust fluid that has more thermal energy than the supply fluid. The arrangement also includes an ejector and a heat exchanger. The ejector is configured to direct at least some of the exhaust fluid into the supply fluid. The heat exchanger is configured to increase thermal energy in the supply fluid using at least some of the exhaust fluid that was not directed into the supply fluid.

Initiation of small-satellite formations via satellite ejector

NASA Astrophysics Data System (ADS)

McMullen, Matthew G

Small satellites can be constructed at a fraction of the cost of a full-size satellite. One full-size satellite can be replaced with a multitude of small satellites, offering expanded area coverage through formation flight. However, the shortcoming to the smaller size is usually a lack of thrusting capabilities. Furthermore, current designs for small satellite deployment mechanisms are only capable of love deployment velocities (on the order of meters per second). Motivated to address this shortcoming, a conceived satellite ejector would offer a significant orbit change by ejecting the satellite at higher deployment velocities (125-200 m/s). Focusing on the applications of the ejector, it is sought to bridge the gap in prior research by offering a method to initiate formation flight. As a precursor to the initiation, the desired orbit properties to initiate the formation are specified in terms of spacing and velocity change vector. From this, a systematic method is developed to find the relationship among velocity change vector, the desired orbit's orientation, and the spacing required to initiate the formation.

Multiplexed operation of a micromachined ultrasonic droplet ejector array.

PubMed

Forbes, Thomas P; Degertekin, F Levent; Fedorov, Andrei G

2007-10-01

A dual-sample ultrasonic droplet ejector array is developed for use as a soft-ionization ion source for multiplexed mass spectrometry (MS). Such a multiplexed ion source aims to reduce MS analysis time for multiple analyte streams, as well as allow for the synchronized ejection of the sample(s) and an internal standard for quantitative results and mass calibration. Multiplexing is achieved at the device level by division of the fluid reservoir and separating the active electrodes of the piezoelectric transducer for isolated application of ultrasonic wave energy to each domain. The transducer is mechanically shaped to further reduce the acoustical crosstalk between the domains. Device design is performed using finite-element analysis simulations and supported by experimental characterization. Isolated ejection of approximately 5 microm diameter water droplets from individual domains in the micromachined droplet ejector array at around 1 MHz frequency is demonstrated by experiments. The proof-of-concept demonstration using a dual-sample device also shows potential for multiplexing with larger numbers of analytes.

Quadrant CFD Analysis of a Mixer-Ejector Nozzle for HSCT Applications

NASA Technical Reports Server (NTRS)

Yoder, Dennis A.; Georgiadis, Nicholas J.; Wolter, John D.

2005-01-01

This study investigates the sidewall effect on flow within the mixing duct downstream of a lobed mixer-ejector nozzle. Simulations which model only one half-chute width of the ejector array are compared with those which model one complete quadrant of the nozzle geometry and with available experimental data. These solutions demonstrate the applicability of the half-chute technique to model the flowfield far away from the sidewall and the necessity of a full-quadrant simulation to predict the formation of a low-energy flow region near the sidewall. The quadrant solutions are further examined to determine the cause of this low-energy region, which reduces the amount of mixing and lowers the thrust of the nozzle. Grid resolution and different grid topologies are also examined. Finally, an assessment of the half-chute and quadrant approaches is made to determine the ability of these simulations to provide qualitative and/or quantitative predictions for this type of complex flowfield.

Multiplexed operation of a micromachined ultrasonic droplet ejector array

DOE Office of Scientific and Technical Information (OSTI.GOV)

Forbes, Thomas P.; Degertekin, F. Levent; Fedorov, Andrei G.

2007-10-15

A dual-sample ultrasonic droplet ejector array is developed for use as a soft-ionization ion source for multiplexed mass spectrometry (MS). Such a multiplexed ion source aims to reduce MS analysis time for multiple analyte streams, as well as allow for the synchronized ejection of the sample(s) and an internal standard for quantitative results and mass calibration. Multiplexing is achieved at the device level by division of the fluid reservoir and separating the active electrodes of the piezoelectric transducer for isolated application of ultrasonic wave energy to each domain. The transducer is mechanically shaped to further reduce the acoustical crosstalk betweenmore » the domains. Device design is performed using finite-element analysis simulations and supported by experimental characterization. Isolated ejection of {approx}5 {mu}m diameter water droplets from individual domains in the micromachined droplet ejector array at around 1 MHz frequency is demonstrated by experiments. The proof-of-concept demonstration using a dual-sample device also shows potential for multiplexing with larger numbers of analytes.« less

Initial development of the two-dimensional ejector shear layer - Experimental results

NASA Technical Reports Server (NTRS)

Benjamin, M. A.; Dufflocq, M.; Roan, V. P.

1993-01-01

An experimental investigation designed to study the development of shear layers in a two-dimensional single-nozzle ejector has been completed. In this study, combinations of air/air, argon/air, helium/air, and air/helium were used as the supersonic primary and subsonic secondary, respectively. Mixing of the gases occurred in a constant-area tube 39.1 mm high by 25.4 mm wide, where the inlet static pressure was maintained at 35 kPa. The cases studied resulted in convective Mach numbers between 0.058 and 1.64, density ratios between 0.102 and 3.49, and velocity ratios between 0.065 and 0.811. The resulting data shows the differences in the shear-layer development for the various combinations of independent variables utilized in the investigation. The normalized growth-rates in the near-field were found to be similar to two-dimensional mixing layers. These results have enhanced the ability to analyze and design ejector systems as well as providing a better understanding of the physics.

Troubleshooting crude vacuum tower overhead ejector systems

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lines, J.R.; Frens, L.L.

1995-03-01

Routinely surveying tower overhead vacuum systems can improve performance and product quality. These vacuum systems normally provide reliable and consistent operation. However, process conditions, supplied utilities, corrosion, erosion and fouling all have an impact on ejector system performance. Refinery vacuum distillation towers use ejector systems to maintain tower top pressure and remove overhead gases. However, as with virtually all refinery equipment, performance may be affected by a number of variables. These variables may act independently or concurrently. It is important to understand basic operating principles of vacuum systems and how performance is affected by: utilities, corrosion and erosion, fouling, andmore » process conditions. Reputable vacuum-system suppliers have service engineers that will come to a refinery to survey the system and troubleshoot performance or offer suggestions for improvement. A skilled vacuum-system engineer may be needed to diagnose and remedy system problems. The affect of these variables on performance is discussed. A case history is described of a vacuum system on a crude tower in a South American refinery.« less

High speed jet noise research at NASA Lewis

NASA Astrophysics Data System (ADS)

Krejsa, Eugene A.; Cooper, B. A.; Kim, C. M.; Khavaran, Abbas

1992-04-01

The source noise portion of the High Speed Research Program at NASA LeRC is focused on jet noise reduction. A number of jet noise reduction concepts are being investigated. These include two concepts, the Pratt & Whitney ejector suppressor nozzle and the General Electric (GE) 2D-CD mixer ejector nozzle, that rely on ejectors to entrain significant amounts of ambient air to mix with the engine exhaust to reduce the final exhaust velocity. Another concept, the GE 'Flade Nozzle' uses fan bypass air at takeoff to reduce the mixed exhaust velocity and to create a fluid shield around a mixer suppressor. Additional concepts are being investigated at Georgia Tech Research Institute and at NASA LeRC. These will be discussed in more detail in later figures. Analytical methods for jet noise prediction are also being developed. Efforts in this area include upgrades to the GE MGB jet mixing noise prediction procedure, evaluation of shock noise prediction procedures, and efforts to predict jet noise directly from the unsteady Navier-Stokes equation.

High speed jet noise research at NASA Lewis

NASA Technical Reports Server (NTRS)

Krejsa, Eugene A.; Cooper, B. A.; Kim, C. M.; Khavaran, Abbas

1992-01-01

The source noise portion of the High Speed Research Program at NASA LeRC is focused on jet noise reduction. A number of jet noise reduction concepts are being investigated. These include two concepts, the Pratt & Whitney ejector suppressor nozzle and the General Electric (GE) 2D-CD mixer ejector nozzle, that rely on ejectors to entrain significant amounts of ambient air to mix with the engine exhaust to reduce the final exhaust velocity. Another concept, the GE 'Flade Nozzle' uses fan bypass air at takeoff to reduce the mixed exhaust velocity and to create a fluid shield around a mixer suppressor. Additional concepts are being investigated at Georgia Tech Research Institute and at NASA LeRC. These will be discussed in more detail in later figures. Analytical methods for jet noise prediction are also being developed. Efforts in this area include upgrades to the GE MGB jet mixing noise prediction procedure, evaluation of shock noise prediction procedures, and efforts to predict jet noise directly from the unsteady Navier-Stokes equation.

Experimental investigation of wing installation effects on a two-dimensional mixer/ejector nozzle for supersonic transport aircraft

NASA Technical Reports Server (NTRS)

Anderson, David J.; Lambert, Heather H.; Mizukami, Masashi

1992-01-01

Experimental results from a wind tunnel test conducted to investigate propulsion/airframe integration (PAI) effects are presented. The objectives of the test were to examine rough order-of-magnitude changes in the acoustic characteristics of a mixer/ejector nozzle due to the presence of a wing and to obtain limited wing and nozzle flow-field measurements. A simple representative supersonic transport wing planform, with deflecting flaps, was installed above a two-dimensional mixer/ejector nozzle that was supplied with high-pressure heated air. Various configurations and wing positions with respect to the nozzle were studied. Because of hardware problems, no acoustics and only a limited set of flow-field data were obtained. For most hardware configurations tested, no significant propulsion/airframe integration effects were identified. Significant effects were seen for extreme flap deflections. The combination of the exploratory nature of the test and the limited flow-field instrumentation made it impossible to identify definitive propulsion/airframe integration effects.

Mean Flow Augmented Acoustics in Rocket Systems

NASA Technical Reports Server (NTRS)

Fischbach, Sean

2014-01-01

Combustion instability in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. Recent advances in energy based modeling of combustion instabilities require accurate determination of acoustic frequencies and mode shapes. Of particular interest is the acoustic mean flow interactions within the converging section of a rocket nozzle, where gradients of pressure, density, and velocity become large. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The present study aims to implement the French model within the COMSOL Multiphysiscs framework and analyzes one of the author's presented test cases.

Low frequency electric field variations during HF transmissions on a mother-daughter rocket

NASA Technical Reports Server (NTRS)

Rosenberg, T. J.; Maynard, M. C.; Holtet, J. A.; Karlsen, N. O.; Egeland, A.; Moe, T. E.; Troim, J.

1977-01-01

HF wave propagation experiments were conducted on Mother-Daughter rockets in the polar ionosphere. Swept frequency transmissions from the Mother, nominally covering the range from 0.5 to 5 MHz in both CW and pulse modes, are received by the Daughter. In the most recent rocket of the series, the Mother also contained an AC electric field spectrometer covering the frequency range from 10 Hz to 100 kHz in four decade bands. The low frequency response of the ionosphere with respect to waves emitted from the onboard HF transmitter is examined.

Tidal analysis of Met rocket wind data

NASA Technical Reports Server (NTRS)

Bedinger, J. F.; Constantinides, E.

1976-01-01

A method of analyzing Met Rocket wind data is described. Modern tidal theory and specialized analytical techniques were used to resolve specific tidal modes and prevailing components in observed wind data. A representation of the wind which is continuous in both space and time was formulated. Such a representation allows direct comparison with theory, allows the derivation of other quantities such as temperature and pressure which in turn may be compared with observed values, and allows the formation of a wind model which extends over a broader range of space and time. Significant diurnal tidal modes with wavelengths of 10 and 7 km were present in the data and were resolved by the analytical technique.

A Method for Prevention of Screaming in Rocket Engines

NASA Technical Reports Server (NTRS)

Kerslake, W. R.; Male, T.

1954-01-01

Lateral and longitudinal combustion-pressure oscillations that occurred in screaming combustion of a 1000-pound-thrust rocket engine using white fuming nitric acid and JP-4 fuel as propellants were successfully prevented by means of longitudinal fins in the combustion chamber. Fin position was critical, and complete attenuation was achieved only when the fins were located in a zone approximately 8 to 16 inches from the injector. Fins located in other zones, that is, near the injector or far downstream from the injector, did not stop the oscillations. When oscillations occurred in finned chambers, the longitudinal mode seemed more dominant than the lateral mode; in chambers without fins, the lateral mode tended to be dominant. The lateral oscillation was distorted and its intensity diminished by the fins. Fins, however, did not affect the frequencies; the longitudinal frequency varied inversely with chamber length, and lateral frequencies varied only slightly from an average of 6000 cycles per second.

Failure mode analysis to predict product reliability.

NASA Technical Reports Server (NTRS)

Zemanick, P. P.

1972-01-01

The failure mode analysis (FMA) is described as a design tool to predict and improve product reliability. The objectives of the failure mode analysis are presented as they influence component design, configuration selection, the product test program, the quality assurance plan, and engineering analysis priorities. The detailed mechanics of performing a failure mode analysis are discussed, including one suggested format. Some practical difficulties of implementation are indicated, drawn from experience with preparing FMAs on the nuclear rocket engine program.

High performance forward swept wing aircraft

NASA Technical Reports Server (NTRS)

Koenig, David G. (Inventor); Aoyagi, Kiyoshi (Inventor); Dudley, Michael R. (Inventor); Schmidt, Susan B. (Inventor)

1988-01-01

A high performance aircraft capable of subsonic, transonic and supersonic speeds employs a forward swept wing planform and at least one first and second solution ejector located on the inboard section of the wing. A high degree of flow control on the inboard sections of the wing is achieved along with improved maneuverability and control of pitch, roll and yaw. Lift loss is delayed to higher angles of attack than in conventional aircraft. In one embodiment the ejectors may be advantageously positioned spanwise on the wing while the ductwork is kept to a minimum.

Calculation of the mixing chamber of an ejector chemical oxygen - iodine laser

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zagidullin, M V; Nikolaev, V D

2001-06-30

Gas parameters are calculated at the outlet of the mixing chamber of an ejector chemical oxygen-iodine laser with a nozzle unit consisting of nozzles of three types, which provides a total pressure of the active medium that substantially exceeds a pressure in the generator of singlet oxygen. This technique of forming the laser active medium substantially facilitates the ejection of the exhaust gas to the atmosphere by using a diffuser and single-stage vacuum systems based on water circulating pumps. (lasers, active media)

Investigation at supersonic and subsonic Mach numbers of auxiliary inlets supplying secondary air flow to ejector exhaust nozzles

NASA Technical Reports Server (NTRS)

Hearth, Donald P; Cubbison, Robert W

1956-01-01

The results indicated increases in auxiliary-inlet pressure recovery with increases in scoop height relative to the boundary-layer thickness. The pressure recovery increased at about the same rate as theoretically predicted for an inlet in a boundary layer having a one-seventh power profile, but was only about 0.68 to 0.75 of the theoretically obtainable values. Under some operating conditions, flow from the primary jet was exhausted through the auxiliary inlet. This phenomenon could be predicted from the ejector pumping characteristics.

Low Cost Hypermixing Ejector Ramjet Program.

DTIC Science & Technology

1975-06-01

SEC.A- SCALE: ZO/I q.zPL vEw A -3 SCALE..?0/1 - w aC LIT OFM TIEA -LLIS S A?-. 87 SECTION XI MODIFIED HYPERMIXING EJECTOR EXPERIMENTAL PROGRAM 1...CAN). psi.. ramage’. Witt shbiuma, .as slacawim iam Fit.,7 Ti’LDtlase lreaa ftime ,mmeasss9ri’d vaiaac’s are aimemit UK ) ft Sv ieAwe(’r 01311 this...Neav OrdueNKw TeOt Station, Dover, New Jersey , March 31, 1955 (Coniderial). 3. Data on UnsyawwetricaI.DinletAyhydrazin* (DMW), Revisin No. 3. Aerojet

SST Technology Follow-On Program - Phase 2. Noise Suppressor/Nozzle Development. Volume 9. Performance Technology - Analysis of the Low Speed Performance of Multitube Suppressor/Ejector Nozzles (0-167 kn)

DTIC Science & Technology

1975-03-01

Layer Suction 18 Temperature and Pressure Profile at Charging Station |9 Roiind-Corivergent Reference Nozzle 20 Elliptical Ramps 21 37-Tube...between plumes of the jets in the outer row of a suppressor Homulary layer Discharge coelticient, accounting for temperature induced no/./Ie area...tunnel floor. The suppressor air tlow rate was measured with an A.S.M.H. long-radius flow nozzle. The boundary layer ihickness at the ejector inlet

Plasma torch testing for thermostructural evaluation of rocket motor nozzle materials

NASA Technical Reports Server (NTRS)

Prince, Andrew S.; Bunker, Robert C.; Lawrence, Tim

1989-01-01

This paper presents data from the thermostructural testing of tape-wrapped carbon phenolic. This work has been performed with the use of a plasma torch and loading device in an effort to study the anomalous erosion characteristicfs of that seen in the Space Shuttle Solid Rocket Motor Nozzle STS-8A. Testing is conducted in an effort to determine conditions or parameters involved in this mode of failure.

A demonstration of an intelligent control system for a reusable rocket engine

NASA Technical Reports Server (NTRS)

Musgrave, Jeffrey L.; Paxson, Daniel E.; Litt, Jonathan S.; Merrill, Walter C.

1992-01-01

An Intelligent Control System for reusable rocket engines is under development at NASA Lewis Research Center. The primary objective is to extend the useful life of a reusable rocket propulsion system while minimizing between flight maintenance and maximizing engine life and performance through improved control and monitoring algorithms and additional sensing and actuation. This paper describes current progress towards proof-of-concept of an Intelligent Control System for the Space Shuttle Main Engine. A subset of identifiable and accommodatable engine failure modes is selected for preliminary demonstration. Failure models are developed retaining only first order effects and included in a simplified nonlinear simulation of the rocket engine for analysis under closed loop control. The engine level coordinator acts as an interface between the diagnostic and control systems, and translates thrust and mixture ratio commands dictated by mission requirements, and engine status (health) into engine operational strategies carried out by a multivariable control. Control reconfiguration achieves fault tolerance if the nominal (healthy engine) control cannot. Each of the aforementioned functionalities is discussed in the context of an example to illustrate the operation of the system in the context of a representative failure. A graphical user interface allows the researcher to monitor the Intelligent Control System and engine performance under various failure modes selected for demonstration.

The Loci Multidisciplinary Simulation System Overview and Status

NASA Technical Reports Server (NTRS)

Luke, Edward A.; Tong, Xiao-Ling; Tang, Lin

2002-01-01

This paper will discuss the Loci system, an innovative tool for developing tightly coupled multidisciplinary three dimensional simulations. This presentation will overview some of the unique capabilities of the Loci system to automate the assembly of numerical simulations from libraries of fundamental computational components. We will discuss the demonstration of the Loci system on coupled fluid-structure problems related to RBCC propulsion systems.

Acoustic-Structure Interaction in Rocket Engines: Validation Testing

NASA Technical Reports Server (NTRS)

Davis, R. Benjamin; Joji, Scott S.; Parks, Russel A.; Brown, Andrew M.

2009-01-01

While analyzing a rocket engine component, it is often necessary to account for any effects that adjacent fluids (e.g., liquid fuels or oxidizers) might have on the structural dynamics of the component. To better characterize the fully coupled fluid-structure system responses, an analytical approach that models the system as a coupled expansion of rigid wall acoustic modes and in vacuo structural modes has been proposed. The present work seeks to experimentally validate this approach. To experimentally observe well-coupled system modes, the test article and fluid cavities are designed such that the uncoupled structural frequencies are comparable to the uncoupled acoustic frequencies. The test measures the natural frequencies, mode shapes, and forced response of cylindrical test articles in contact with fluid-filled cylindrical and/or annular cavities. The test article is excited with a stinger and the fluid-loaded response is acquired using a laser-doppler vibrometer. The experimentally determined fluid-loaded natural frequencies are compared directly to the results of the analytical model. Due to the geometric configuration of the test article, the analytical model is found to be valid for natural modes with circumferential wave numbers greater than four. In the case of these modes, the natural frequencies predicted by the analytical model demonstrate excellent agreement with the experimentally determined natural frequencies.

Ventilation through a small-bore catheter: optimizing expiratory ventilation assistance.

PubMed

Hamaekers, A E W; Borg, P A J; Götz, T; Enk, D

2011-03-01

Emergency ventilation through a small-bore transtracheal catheter can be lifesaving in a 'cannot intubate, cannot ventilate' situation. Ejectors, capable of creating suction by the Bernoulli principle, have been proposed to facilitate expiration through small-bore catheters. In this bench study, we compared a novel, purpose-built ventilation ejector (DE 5) with a previously proposed, modified industrial ejector (SBP 07). The generated insufflation pressures, suction pressures in static and dynamic situations, and also suction capacities and entrainment ratios of the SBP 07 and the DE 5 were determined. The DE 5 was also tested in a lung simulator with a simulated complete upper airway obstruction. Inspiratory and expiratory times through a transtracheal catheter were measured at various flow rates and achievable minute volumes were calculated. In a static situation, the SBP 07 showed a more negative pressure build-up compared with the DE 5. However, in a dynamic situation, the DE 5 generated a more negative pressure, resulting in a higher suction capacity. Employment of the DE 5 at a flow rate of 18 litre min(-1) allowed a minute volume through the transtracheal catheter of up to 8.27 litre min(-1) at a compliance of 100 ml cm H(2)O(-1). The efficiency of the DE 5 depended on the flow rate of the driving gas and the compliance of the lung simulator. In laboratory tests, the DE 5 is an optimized ventilation ejector suitable for applying expiratory ventilation assistance. Further research may confirm the clinical applicability as a portable emergency ventilator for use with small-bore catheters.

Water impact test of aft skirt end ring, and mid ring segments of the Space Shuttle Solid Rocket Booster

NASA Technical Reports Server (NTRS)

1983-01-01

The results of water impact loads tests using aft skirt end ring, and mid ring segments of the Space Shuttle Solid Rocket Booster (SRB) are examined. Dynamic structural response data is developed and an evaluation of the model in various configurations is presented. Impact velocities are determined for the SRB with the larger main chute system. Various failure modes are also investigated.

Rocket Chamber Temperature Measurements by Microwave Techniques

DTIC Science & Technology

1974-07-01

acoustic oscillation inside a cylindrical end burner la theoretically derived and experimentally observed. It.« oscillation frequencies observed range...from 3.2 to 4.4 kHz, whereas the theoretic?! oscillation frequencies range from 2.98 to 5.13 kHz for various oscillation modes. Acoustic gain and...loss expressions are derived and applied to the rocket firings. The results show that for a atable system, the acoustic loss exceed« the acoustic

A mathematical model for simulating noise suppression of lined ejectors

NASA Technical Reports Server (NTRS)

Watson, Willie R.

1994-01-01

A mathematical model containing the essential features embodied in the noise suppression of lined ejectors is presented. Although some simplification of the physics is necessary to render the model mathematically tractable, the current model is the most versatile and technologically advanced at the current time. A system of linearized equations and the boundary conditions governing the sound field are derived starting from the equations of fluid dynamics. A nonreflecting boundary condition is developed. In view of the complex nature of the equations, a parametric study requires the use of numerical techniques and modern computers. A finite element algorithm that solves the differential equations coupled with the boundary condition is then introduced. The numerical method results in a matrix equation with several hundred thousand degrees of freedom that is solved efficiently on a supercomputer. The model is validated by comparing results either with exact solutions or with approximate solutions from other works. In each case, excellent correlations are obtained. The usefulness of the model as an optimization tool and the importance of variable impedance liners as a mechanism for achieving broadband suppression within a lined ejector are demonstrated.

Acoustic Characteristics of Various Treatment Panel Designs Specific to HSCT Mixer-Ejector Application

NASA Technical Reports Server (NTRS)

Salikuddin, M.; Kinzie, K.; Vu, D. D.; Langenbrunner, L. E.; Szczepkowski, G. T.

2006-01-01

The development process of liner design methodology is described in several reports. The results of the initial effort of concept development, screening, laboratory testing of various liner concepts, and preliminary correlation (generic data) are presented in a report Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program. The second phase of laboratory test results of more practical concepts and their data correlations are presented in this report (product specific). In particular, this report contains normal incidence impedance measurements of several liner types in both a static rig and in a high temperature flow duct rig. The flow duct rig allows for temperatures up to 400 F with a grazing flow up to Mach 0.8. Measurements of impedance, DC flow resistance, and in the flow rig cases, impact of the liner on boundary layer profiles are documented. In addition to liner rig tests, a limited number of tests were made on liners installed in a mixer-Ejector nozzle to confirm the performance of the liner prediction in an installed configuration.

Assessment of Integrated Nozzle Performance

NASA Technical Reports Server (NTRS)

Lambert, H. H.; Mizukami, M.

1999-01-01

This presentation highlights the activities that researchers at the NASA Lewis Research Center (LeRC) have been and will be involved in to assess integrated nozzle performance. Three different test activities are discussed. First, the results of the Propulsion Airframe Integration for High Speed Research 1 (PAIHSR1) study are presented. The PAIHSR1 experiment was conducted in the LeRC 9 ft x l5 ft wind tunnel from December 1991 to January 1992. Second, an overview of the proposed Mixer/ejector Inlet Distortion Study (MIDIS-E) is presented. The objective of MIDIS-E is to assess the effects of applying discrete disturbances to the ejector inlet flow on the acoustic and aero-performance of a mixer/ejector nozzle. Finally, an overview of the High-Lift Engine Aero-acoustic Technology (HEAT) test is presented. The HEAT test is a cooperative effort between the propulsion system and high-lift device research communities to assess wing/nozzle integration effects. The experiment is scheduled for FY94 in the NASA Ames Research Center (ARC) 40 ft x 80 ft Low Speed Wind Tunnel (LSWT).

Spray combustion under oscillatory pressure conditions

NASA Technical Reports Server (NTRS)

Jacobs, H. R.; Santoro, R. J.

1991-01-01

The performance and stability of liquid rocket engines is often argued to be significantly impacted by atomization and droplet vaporization processes. In particular, combustion instability phenomena may result from the interactions between the oscillating pressure field present in the rocket combustor and the fuel and oxidizer injection process. Few studies have been conducted to examine the effects of oscillating pressure fields on spray formation and its evolution under rocket engine conditions. The pressure study is intended to address the need for such studies. In particular, two potentially important phenomena are addressed in the present effort. The first involves the enhancement of the atomization process for a liquid jet subjected to an oscillating pressure field of known frequency and amplitude. The objective of this part of the study is to examine the coupling between the pressure field and or the resulting periodically perturbed velocity field on the breakup of the liquid jet. In particular, transverse mode oscillations are of interest since such modes are considered of primary importance in combustion instability phenomena. The second aspect of the project involves the effects of an oscillating pressure on droplet coagulation and secondary atomization. The objective of this study is to examine the conditions under which phenomena following the atomization process are affected by perturbations to the pressure or velocity fields. Both coagulation and represent a coupling mechanism between the pressure field and the energy release process in rocket combustors. It is precisely this coupling which drives combustion instability phenomena. Consequently, the present effort is intended to provide the fundamental insights needed to evaluate these processes as important mechanisms in liquid rocket instability phenomena.

Rocket in situ observation of equatorial plasma irregularities in the region between E and F layers over Brazil

NASA Astrophysics Data System (ADS)

Savio Odriozola, Siomel; de Meneses, Francisco Carlos, Jr.; Muralikrishna, Polinaya; Alvares Pimenta, Alexandre; Alam Kherani, Esfhan

2017-03-01

A two-stage VS-30 Orion rocket was launched from the equatorial rocket launching station in Alcântara, Brazil, on 8 December 2012 soon after sunset (19:00 LT), carrying a Langmuir probe operating alternately in swept and constant bias modes. At the time of launch, ground equipment operated at equatorial stations showed rapid rise in the base of the F layer, indicating the pre-reversal enhancement of the F region vertical drift and creating ionospheric conditions favorable for the generation of plasma bubbles. Vertical profiles of electron density estimated from Langmuir probe data showed wave patterns and small- and medium-scale plasma irregularities in the valley region (100-300 km) during the rocket upleg and downleg. These irregularities resemble those detected by the very high frequency (VHF) radar installed at Jicamarca and so-called equatorial quasi-periodic echoes. We present evidence suggesting that these observations could be the first detection of this type of irregularity made by instruments onboard a rocket.

Water outlet control mechanism for fuel cell system operation in variable gravity environments

NASA Technical Reports Server (NTRS)

Vasquez, Arturo (Inventor); McCurdy, Kerri L. (Inventor); Bradley, Karla F. (Inventor)

2007-01-01

A self-regulated water separator provides centrifugal separation of fuel cell product water from oxidant gas. The system uses the flow energy of the fuel cell's two-phase water and oxidant flow stream and a regulated ejector or other reactant circulation pump providing the two-phase fluid flow. The system further uses a means of controlling the water outlet flow rate away from the water separator that uses both the ejector's or reactant pump's supply pressure and a compressibility sensor to provide overall control of separated water flow either back to the separator or away from the separator.

A Summary/Overview of Ejector Augmentor Theory and Performance. Volume 1. Technical Discussion

DTIC Science & Technology

1979-09-01

P Tt ’ Pt m m P - tp pp ptm Pt P Pt Tt n CONSTANT PRESSURE MIXIN -,"--"MIXING INITIAL HEAT ADDITION TO THE PRIMARY w It s sm S Sm ENTROPY, s FfGURE 2...TEMPERATURE-ENTROPY DIAGRAM FOR AN EJECTOR CYCLE ........ . .. I.... ...T H NITIAL HEAT ADDITION TO THE PRIMARY s s sm s8 psi Pt s Tt s ex AIR Uu...s Sm = (8) This Increase in entropy drives the cycle performance to lower values of calcu- lated p, as shown in Figure 4. The only boundary condition

Venturi Air-Jet Vacuum Ejector For Sampling Air

NASA Technical Reports Server (NTRS)

Hill, Gerald F.; Sachse, Glen W.; Burney, L. Garland; Wade, Larry O.

1990-01-01

Venturi air-jet vacuum ejector pump light in weight, requires no electrical power, does not contribute heat to aircraft, and provides high pumping speeds at moderate suctions. High-pressure motive gas required for this type of pump bled from compressor of aircraft engine with negligible effect on performance of engine. Used as source of vacuum for differential-absorption CO-measurement (DACOM), modified to achieve in situ measurements of CO at frequency response of 10 Hz. Provides improvement in spatial resolution and potentially leads to capability to measure turbulent flux of CO by use of eddy-correlation technique.

Real-time simulation of an F110/STOVL turbofan engine

NASA Technical Reports Server (NTRS)

Drummond, Colin K.; Ouzts, Peter J.

1989-01-01

A traditional F110-type turbofan engine model was extended to include a ventral nozzle and two thrust-augmenting ejectors for Short Take-Off Vertical Landing (STOVL) aircraft applications. Development of the real-time F110/STOVL simulation required special attention to the modeling approach to component performance maps, the low pressure turbine exit mixing region, and the tailpipe dynamic approximation. Simulation validation derives by comparing output from the ADSIM simulation with the output for a validated F110/STOVL General Electric Aircraft Engines FORTRAN deck. General Electric substantiated basic engine component characteristics through factory testing and full scale ejector data.

Entrainment and mixing in thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Bernal, L.; Sarohia, V.

1983-01-01

An experimental investigation of two-dimensional thrust augmenting ejector flows has been conducted. Measurements of the shroud surface pressure distribution, mean velocity, turbulent intensities and Reynolds stresses were made in two shroud geometries at various primary nozzle pressure ratios. The effects of shroud geometry and primary nozzle pressure ratio on the shroud surface pressure distribution, mean flow field and turbulent field were determined. From these measurements the evolution of mixing within the shroud of the primary flow and entrained fluid was obtained. The relationship between the mean flow field, the turbulent field and the shroud surface pressure distribution is discussed.

Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

NASA Image and Video Library

1999-08-25

Technicians prepare a Pegasus rocket booster for flight tests with the X-43A "Hypersonic Experimental Vehicle," or "Hyper-X." The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude).

Modal Analysis of Space-rocket Equipment Components

NASA Astrophysics Data System (ADS)

Igolkin, A. A.; Safin, A. I.; Prokofiev, A. B.

2018-01-01

In order to prevent vibration damage an analysis of natural frequencies and mode shapes of elements of rocket and space technology should be developed. This paper discusses technique of modal analysis on the example of the carrier platform. Modal analysis was performed by using mathematical modeling and laser vibrometer. Experimental data was clarified by using Test.Lab software. As a result of modal analysis amplitude-frequency response of carrier platform was obtained and the parameters of the elasticity was clarified.

Moving base simulation of an ASTOVL lift-fan aircraft

NASA Technical Reports Server (NTRS)

Chung, William W. Y.; Borchers, Paul F.; Franklin, James A.

1995-01-01

Using a generalized simulation model, a moving-base simulation of a lift-fan short takeoff/vertical landing fighter aircraft was conducted on the Vertical Motion Simulator at Ames Research Center. Objectives of the experiment were to (1) assess the effects of lift-fan propulsion system design features on aircraft control during transition and vertical flight including integration of lift fan/lift/cruise engine/aerodynamic controls and lift fan/lift/cruise engine dynamic response, (2) evaluate pilot-vehicle interface with the control system and head-up display including control modes for low-speed operational tasks and control mode/display integration, and (3) conduct operational evaluations of this configuration during takeoff, transition, and landing similar to those carried out previously by the Ames team for the mixed-flow, vectored thrust, and augmentor-ejector concepts. Based on results of the simulation, preliminary assessments of acceptable and borderline lift-fan and lift/cruise engine thrust response characteristics were obtained. Maximum pitch, roll, and yaw control power used during transition, hover, and vertical landing were documented. Control and display mode options were assessed for their compatibility with a range of land-based and shipboard operations from takeoff to cruise through transition back to hover and vertical landing. Flying qualities were established for candidate control modes and displays for instrument approaches and vertical landings aboard an LPH assault ship and DD-963 destroyer. Test pilot and engineer teams from the Naval Air Warfare Center, Boeing, Lockheed, McDonnell Douglas, and the British Defence Research Agency participated in the program.

Mixing Process in Ejector Nozzles Studied at Lewis' Aero-Acoustic Propulsion Laboratory

NASA Technical Reports Server (NTRS)

1996-01-01

The NASA Lewis Research Center has been studying mixing processes in ejector nozzles for its High Speed Research (HSR) Program. This work is directed at finding ways to minimize the noise of a future supersonic airliner. Much of the noise such an airplane would generate would come from the nozzle, where a hot, high-speed jet exits the engine. Several different nozzle configurations were used to produce nozzle systems with different acoustical and aerodynamic characteristics. The acoustical properties were measured by an array of microphones in an anechoic chamber, and the aerodynamics were measured by traditional pressure and temperature instruments as well as by Laser Doppler Velocimetry (LDV), a technique for visualizing the airflow pattern without disturbing it. These measurements were put together and compared for different configurations to examine the relationships between mixing and noise generation. The mixer-ejector nozzle with the installed flow-visualization windows (foreground), the optical equipment and the supporting structure for the Laser Doppler Velocimetry flow visualization (midfield), and the sound-absorbing wedges used to create an anechoic environment for acoustic testing (background) is shown. The High Speed Research Program is a NASA-funded effort, in cooperation with the U.S. aerospace industry, to develop enabling technologies for a future supersonic airliner. One of the technological barriers being addressed is noise generated during near-airport operation. The mixer-ejector nozzle concept is being examined as a way to reduce jet noise while maintaining thrust. Ambient air is mixed with the high-velocity engine exhaust to reduce the jet velocity and hence the noise generated by the jet. The model was designed and built by Pratt & Whitney under NASA contract. The test, completed in June 1995, was conducted in Lewis' Aero-Acoustic Propulsion Laboratory.

Plate Deformation Behavior of Polymer Matrix Composite-Ti Honeycomb-Metal Sandwiches for Pressurized Propulsion Component Applications

NASA Technical Reports Server (NTRS)

Bertelsen, William D.; Shin, E. eugene; Thesken, John C.; Sutter, James K.; Martin, Rich

2004-01-01

THe objectives are: 1. To experimentally validate bi-axial plate flexural performance of PMC-Ti H/C-A286 sandwich panels for the internally pressurized RBCC combustion chamber support structure. 2. To explore ASTM 2-D plate flexure test (D 6416) to simulate the internal pressure loading and to correlate the results with analytical and FE modeling based on 2-D flexure properties.

Lung function and blood markers of nutritional status in non-COPD aging men with smoking history: a cross-sectional study.

PubMed

Shiozawa, Nobuyoshi; Hayashimoto, Kanae; Suzuki, Etsuji; Kikuchi, Hiroshi; Takata, Shingo; Ashida, Kozo; Watanabe, Masutaka; Hosaki, Yasuhiro; Mitsunobu, Fumihiro

2010-08-09

Cigarette smoking and advanced age are well known as risk factors for chronic obstructive pulmonary disease (COPD), and nutritional abnormalities are important in patients with COPD. However, little is known about the nutritional status in non-COPD aging men with smoking history. We therefore investigated whether reduced lung function is associated with lower blood markers of nutritional status in those men. This association was examined in a cross-sectional study of 65 Japanese male current or former smokers aged 50 to 80 years: 48 without COPD (non-COPD group), divided into tertiles according to forced expiratory volume in one second as percent of forced vital capacity (FEV(1)/FVC), and 17 with COPD (COPD group). After adjustment for potential confounders, lower FEV(1)/FVC was significantly associated with lower red blood cell count (RBCc), hemoglobin, and total protein (TP); not with total energy intake. The difference in adjusted RBCc and TP among the non-COPD group tertiles was greater than that between the bottom tertile in the non-COPD group and the COPD group. In non-COPD aging men with smoking history, trends toward reduced nutritional status and anemia may independently emerge in blood components along with decreased lung function even before COPD onset.

Studies of the differential absorption rocket experiment. [to measure atmospheric electron density

NASA Technical Reports Server (NTRS)

Ginther, J. C.; Smith, L. G.

1975-01-01

Investigations of the ionosphere, in the rocket program of the Aeronomy Laboratory, include a propagation experiment, the data from which may be analyzed in several modes. This report considers in detail the differential absorption experiment. The sources of error and limitations of sensitivity are discussed. Methods of enhancing the performance of the experiment are described. Some changes have been made in the system and the improvement demonstrated. Suggestions are made for further development of the experiment.

Mixing in Shear Coaxial Jets with and without Acoustics (Briefing Charts)

DTIC Science & Technology

2012-05-21

and heat transfer fluctuations in a rocket engine – Irreparable damage can occur in əs • Combustion Instability caused a 4-yr delay in the...common choice for cryogenic liquid rocket engines • Interactions of transverse acoustics with injector’s own modes and mixing needs to be understood...Pr = 0.44 • LAR-thin , Pr = 0.44, J = 0.5 POM 2 POM 1 Average Snapshot Power Spectral Densities (PSD) of Temporal Coefficients of POMs 1 and 2

Combustion dynamics in liquid rocket engines

NASA Technical Reports Server (NTRS)

Mclain, W. H.

1971-01-01

A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.

Efficacy of the ejector flow-meter. A scavenging device for anaesthetic gases.

PubMed

Obel, D; Jørgensen, S; Ferguson, A; Frandsen, K

1985-01-01

Measurements of air concentrations of nitrous oxide and halothane in the breathing zone of the anaesthetist and the operating-room nurse were carried out during inhalation anaesthesia with a Mapleson D system. Gas removal was performed from inside the breathing system at the same rate as that of the fresh gas inflow by means of an ejector flow-meter. The concentrations of nitrous oxide and halothane were maintained below the Danish Threshold Limit Values of 100 and 5 parts per million, respectively, by using this type of scavenging. When these anaesthetics were used simultaneously, the reduced Threshold Limit Values were not exceeded during endotracheal anaesthesia.

Vacuum generation in pneumatic artificial heart drives with a specially designed ejector system.

PubMed

Schima, H; Huber, L; Spitaler, F

1990-06-01

To improve the filling characteristics of pneumatically driven membrane artificial hearts (AHs), a vacuum is applied during diastole. This paper describes an ejector system for AH-drivers based on the Venturi effect, which was designed for this purpose. It provides vacuums of more than -40 mmHg at flow rates up to 50 l/min requiring a supplying primary gas pressure of less than 150 kPa (1140 mmHg). Under normal working conditions, the necessary supply flow was less than 5l/min. The device is small, cheap, quiet and fail-safe, and has been evaluated successfully in experimental and clinical use.

Delivering supplemental oxygen during sedation via a saliva ejector.

PubMed

Milnes, Alan R

2002-01-01

Intraoperative oxygen supplementation to sedated children has been shown to prevent hemoglobin desaturations even in the presence of apnea during pediatric conscious sedation. Although many practitioners deliver supplemental oxygen via a nasal hood, this method is impractical and often unsuccessful if the child is a mouth breather, has moderate adenotonsillar hypertrophy or occasionally cries during treatment (at which time there will be mouth breathing). This paper describes a method in which the saliva ejector is used to deliver supplemental oxygen to sedated children while they are receiving dental treatment. The advantages of this method and suggestions for its successful application are also included.

A simple apparatus for the experimental study of non-steady flow thrust-augmenter ejector configurations

NASA Technical Reports Server (NTRS)

Khare, J. M.; Kentfield, J. A. C.

1979-01-01

A flexible, and easily modified, test rig is described which allows a one dimensional nonsteady flow stream to be generated, economically from a steady flow source of compressed air. This nonsteady flow is used as the primary stream in a nonsteady flow ejector constituting part of the test equipment. Standard piezo-electric pressure transducers etc. allow local pressures to be studied, as functions of time, in both the primary and secondary (mixed) flow portions of the apparatus. Provision is also made for measuring the primary and secondary mass flows and the thrust generated. Sample results obtained with the equipment are presented.

Performance comparison of supersonic ejectors with different motive gas injection schemes applicable for flowing medium gas laser

NASA Astrophysics Data System (ADS)

Singhal, G.; Subbarao, P. M. V.; Mainuddin; Tyagi, R. K.; Dawar, A. L.

2017-05-01

A class of flowing medium gas lasers with low generator pressures employ supersonic flows with low cavity pressure and are primarily categorized as high throughput systems capable of being scaled up to MW class. These include; Chemical Oxygen Iodine Laser (COIL) and Hydrogen (Deuterium) Fluoride (HF/DF). The practicability of such laser systems for various applications is enhanced by exhausting the effluents directly to ambient atmosphere. Consequently, ejector based pressure recovery forms a potent configuration for open cycle operation. Conventionally these gas laser systems require at least two ejector stages with low pressure stage being more critical, since it directly entrains the laser media, and the ensuing perturbation of cavity flow, if any, may affect laser operation. Hence, the choice of plausible motive gas injection schemes viz., peripheral or central is a fluid dynamic issue of interest, and a parametric experimental performance comparison would be beneficial. Thus, the focus is to experimentally characterize the effect of variation in motive gas supply pressure, entrainment ratio, back pressure conditions, nozzle injection position operated together with a COIL device and discern the reasons for the behavior.

Analytical and computational studies on the vacuum performance of a chevron ejector

NASA Astrophysics Data System (ADS)

Kong, F. S.; Jin, Y. Z.; Kim, H. D.

2016-11-01

The effects of chevrons on the performance of a supersonic vacuum ejector-diffuser system are investigated numerically and evaluated theoretically in this work. A three-dimensional geometrical domain is numerically solved using a fully implicit finite volume scheme based on the unsteady Reynolds stress model. A one-dimensional mathematical model provides a useful tool to reveal the steady flow physics inside the vacuum ejector-diffuser system. The effects of the chevron nozzle on the generation of recirculation regions and Reynolds stress behaviors are studied and compared with those of a conventional convergent nozzle. The present performance parameters obtained from the simulated results and the mathematical results are validated with existing experimental data and show good agreement. Primary results show that the duration of the transient period and the secondary chamber pressure at a dynamic equilibrium state depend strongly on the primary jet conditions, such as inlet pressure and primary nozzle shape. Complicated oscillatory flow, generated by the unsteady movement of recirculation, finally settles into a dynamic equilibrium state. As a vortex generator, the chevron demonstrated its strong entrainment capacity to accelerate the starting transient flows to a certain extent and reduce the dynamic equilibrium pressure of the secondary chamber significantly.

Chembio extraction on a chip by nanoliter droplet ejection.

PubMed

Yu, Hongyu; Kwon, Jae Wan; Kim, Eun Sok

2005-03-01

This paper describes a novel liquid separation technique for chembio extraction by an ultrasonic nanoliter-liquid-droplet ejector built on a PZT sheet. This technique extracts material from an aqueous two-phase system (ATPS) in a precise amount through digital control of the number of nanoliter droplets, without any mixing between the two liquids in the ATPS. The ultrasonic droplet ejector uses an acoustic streaming effect produced by an acoustic beam focused on the liquid surface, and ejects liquid droplets only from the liquid surface without disturbing most of the liquid below the surface. This unique characteristic of the focused acoustic beam is perfect (1) for separating a top-layer liquid (from the bulk of liquid) that contains particles of interest or (2) for recovering a top-layer liquid that has different phase from a bottom-layer liquid. Three kinds of liquid extraction are demonstrated with the ultrasonic droplet ejector: (1) 16 microl of top layer in Dextran-polyethylene glycol-water ATPS (aqueous two-phase system) is recovered within 20 s; (2) micron sized particles that float on water surface are ejected out with water droplets; and (3) oil layer on top of water is separated out.

Improving turbine performance by cooling inlet air using a waste heat powered ejector refrigerator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kowalski, G.J.

1996-12-31

Stationary turbines are used to produce electricity in many areas of the world. Their performance is adversely affected by high ambient temperatures. Several means of reducing the turbine inlet temperature (offpeak water chiller and ice storage and absorption refrigeration systems) are being proposed as a means of increasing turbine output. In the present investigation the feasibility of increasing turbine output power by using its exhaust gases to power an ejector refrigeration system is demonstrated. The advantages of the ejector refrigeration are: it operates on a non-CFC fluid, its small number of moving parts and its small size. The analysis focusesmore » on United Technologies FT4 turbine with a base load output of 21.6 MW. It is demonstrated that the proposed system can decrease the turbine inlet temperature from 296.2 K to 277.6 K which increases the turbine output by 12.8% during periods of high ambient temperature and improves yearly averaged power output by 5.5% in a temperature climate. It is shown that the energy in the turbine exhaust has the potential of producing additional cooling beyond that required to reduce the inlet temperature.« less

A full-scale STOVL ejector experiment

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy S.

1993-01-01

The design and development of thrust augmenting short take-off and vertical landing (STOVL) ejectors has typically been an iterative process. In this investigation, static performance tests of a full-scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators, yarn tufts and paint dots) was used to assess inlet flowfield characteristics, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate 'seasonal' aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature. Full-scale experimental tests such as this are expensive, and difficult to implement at engine exhaust temperatures. For this reason the utility of using similarity principles -- in particular, the Munk and Prim similarity principle for isentropic flow -- was explored. At different primary temperatures, exit pressure contours are compared for similarity. A nondimensional flow parameter is then shown to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments. Under the assumption that an appropriate similarity principle can be established, then properly chosen performance parameters should be similar for both hot flow and cold flow model tests.

On the scaling of small-scale jet noise to large scale

NASA Technical Reports Server (NTRS)

Soderman, Paul T.; Allen, Christopher S.

1992-01-01

An examination was made of several published jet noise studies for the purpose of evaluating scale effects important to the simulation of jet aeroacoustics. Several studies confirmed that small conical jets, one as small as 59 mm diameter, could be used to correctly simulate the overall or perceived noise level (PNL) noise of large jets dominated by mixing noise. However, the detailed acoustic spectra of large jets are more difficult to simulate because of the lack of broad-band turbulence spectra in small jets. One study indicated that a jet Reynolds number of 5 x 10(exp 6) based on exhaust diameter enabled the generation of broad-band noise representative of large jet mixing noise. Jet suppressor aeroacoustics is even more difficult to simulate at small scale because of the small mixer nozzles with flows sensitive to Reynolds number. Likewise, one study showed incorrect ejector mixing and entrainment using a small-scale, short ejector that led to poor acoustic scaling. Conversely, fairly good results were found with a longer ejector and, in a different study, with a 32-chute suppressor nozzle. Finally, it was found that small-scale aeroacoustic resonance produced by jets impacting ground boards does not reproduce at large scale.

On the scaling of small-scale jet noise to large scale

NASA Technical Reports Server (NTRS)

Soderman, Paul T.; Allen, Christopher S.

1992-01-01

An examination was made of several published jet noise studies for the purpose of evaluating scale effects important to the simulation of jet aeroacoustics. Several studies confirmed that small conical jets, one as small as 59 mm diameter, could be used to correctly simulate the overall or PNL noise of large jets dominated by mixing noise. However, the detailed acoustic spectra of large jets are more difficult to simulate because of the lack of broad-band turbulence spectra in small jets. One study indicated that a jet Reynolds number of 5 x 10 exp 6 based on exhaust diameter enabled the generation of broad-band noise representative of large jet mixing noise. Jet suppressor aeroacoustics is even more difficult to simulate at small scale because of the small mixer nozzles with flows sensitive to Reynolds number. Likewise, one study showed incorrect ejector mixing and entrainment using small-scale, short ejector that led to poor acoustic scaling. Conversely, fairly good results were found with a longer ejector and, in a different study, with a 32-chute suppressor nozzle. Finally, it was found that small-scale aeroacoustic resonance produced by jets impacting ground boards does not reproduce at large scale.

Credit BG. View looking northeast down from the tower onto ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Credit BG. View looking northeast down from the tower onto the two horizontal test stations at Test Stand "D." Station Dy is at the far left (Dy vacuum cell out of view), with in-line exhaust gas cooling sections and steam-driven "air ejector" (or evacuator) discharging engine exhausts to the east. The Dd cell is visible at the lower left, and the Dd exhaust train has the same functions as at Dy. The spherical tank is an electrically heated "accumulator" which supplies steam to the ejectors at Dv, Dd, and Dy stations. Other large piping delivered cooling water to the horizontal train cooling sections. The horizontal duct at the "Y" branch in the Dd train connects the Dd ejector to the Dv and Cv vacuum duct system (a blank can be bolted into this duct to isolate the Dd system). The shed roof for the Dpond test station appears at bottom center of this image. The open steel frame to the lower left of the image supports a hoist and crane for installing or removing test engines from the Dd test cell - Jet Propulsion Laboratory Edwards Facility, Test Stand D, Edwards Air Force Base, Boron, Kern County, CA

Nozzle Aerodynamic Stability During a Throat Shift

NASA Technical Reports Server (NTRS)

Kawecki, Edwin J.; Ribeiro, Gregg L.

2005-01-01

An experimental investigation was conducted on the internal aerodynamic stability of a family of two-dimensional (2-D) High Speed Civil Transport (HSCT) nozzle concepts. These nozzles function during takeoff as mixer-ejectors to meet acoustic requirements, and then convert to conventional high-performance convergent-divergent (CD) nozzles at cruise. The transition between takeoff mode and cruise mode results in the aerodynamic throat and the minimum cross-sectional area that controls the engine backpressure shifting location within the nozzle. The stability and steadiness of the nozzle aerodynamics during this so called throat shift process can directly affect the engine aerodynamic stability, and the mechanical design of the nozzle. The objective of the study was to determine if pressure spikes or other perturbations occurred during the throat shift process and, if so, identify the caused mechanisms for the perturbations. The two nozzle concepts modeled in the test program were the fixed chute (FC) and downstream mixer (DSM). These 2-D nozzles differ principally in that the FC has a large over-area between the forward throat and aft throat locations, while the DSM has an over-area of only about 10 percent. The conclusions were that engine mass flow and backpressure can be held constant simultaneously during nozzle throat shifts on this class of nozzles, and mode shifts can be accomplished at a constant mass flow and engine backpressure without upstream pressure perturbations.

Solid Propellant Grain Structural Integrity Analysis

NASA Technical Reports Server (NTRS)

1973-01-01

The structural properties of solid propellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.

Determination of the availability of appropriate aged flight rocket motors. [captive tests to determine case bond separation and grain bore cracking

NASA Technical Reports Server (NTRS)

Martin, P. J.

1974-01-01

A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

A Flight Demonstration of Plasma Rocket Propulsion

NASA Technical Reports Server (NTRS)

Petro, Andrew; Chang-Diaz, Franklin; Schwenterly, WIlliam; Hitt, Michael; Lepore, Joseph

2000-01-01

The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a variable specific impulse magnetoplasma rocket (V ASIMR) for several years. This type of rocket could be used in the future to propel interplanetary spacecraft and has the potential to open the entire solar system to human exploration. One feature of this propulsion technology is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. Variation of specific impulse and thrust enhances the ability to optimize interplanetary trajectories and results in shorter trip times and lower propellant requirements than with a fixed specific impulse. In its ultimate application for interplanetary travel, the VASIMR would be a multi-megawatt device. A much lower power system is being designed for demonstration in the 2004 timeframe. This first space demonstration would employ a lO-kilowatt thruster aboard a solar powered spacecraft in Earth orbit. The 1O-kilowatt V ASIMR demonstration unit would operate for a period of several months with hydrogen or deuterium propellant with a specific impulse of 10,000 seconds.

This Is Rocket Science!

NASA Astrophysics Data System (ADS)

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-09-01

Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

Droplet Core Nuclear Rocket (DCNR)

NASA Technical Reports Server (NTRS)

Anghaie, Samim

1991-01-01

The most basic design feature of the droplet core nuclear reactor is to spray liquid uranium into the core in the form of droplets on the order of five to ten microns in size, to bring the reactor to critical conditions. The liquid uranium fuel ejector is driven by hydrogen, and more hydrogen is injected from the side of the reactor to about one and a half meters from the top. High temperature hydrogen is expanded through a nozzle to produce thrust. The hydrogen pressure in the system can be somewhere between 50 and 500 atmospheres; the higher pressure is more desirable. In the lower core region, hydrogen is tangentially injected to serve two purposes: (1) to provide a swirling flow to protect the wall from impingement of hot uranium droplets: (2) to generate a vortex flow that can be used for fuel separation. The reactor is designed to maximize the energy generation in the upper region of the core. The system can result in and Isp of 2000 per second, and a thrust-to-weight ratio of 1.6 for the shielded reactor. The nuclear engine system can reduce the Mars mission duration to less than 200 days. It can reduce the hydrogen consumption by a factor of 2 to 3, which reduces the hydrogen load by about 130 to 150 metric tons.

Real-time diagnostics for a reusable rocket engine

NASA Technical Reports Server (NTRS)

Guo, T. H.; Merrill, W.; Duyar, A.

1992-01-01

A hierarchical, decentralized diagnostic system is proposed for the Real-Time Diagnostic System component of the Intelligent Control System (ICS) for reusable rocket engines. The proposed diagnostic system has three layers of information processing: condition monitoring, fault mode detection, and expert system diagnostics. The condition monitoring layer is the first level of signal processing. Here, important features of the sensor data are extracted. These processed data are then used by the higher level fault mode detection layer to do preliminary diagnosis on potential faults at the component level. Because of the closely coupled nature of the rocket engine propulsion system components, it is expected that a given engine condition may trigger more than one fault mode detector. Expert knowledge is needed to resolve the conflicting reports from the various failure mode detectors. This is the function of the diagnostic expert layer. Here, the heuristic nature of this decision process makes it desirable to use an expert system approach. Implementation of the real-time diagnostic system described above requires a wide spectrum of information processing capability. Generally, in the condition monitoring layer, fast data processing is often needed for feature extraction and signal conditioning. This is usually followed by some detection logic to determine the selected faults on the component level. Three different techniques are used to attack different fault detection problems in the NASA LeRC ICS testbed simulation. The first technique employed is the neural network application for real-time sensor validation which includes failure detection, isolation, and accommodation. The second approach demonstrated is the model-based fault diagnosis system using on-line parameter identification. Besides these model based diagnostic schemes, there are still many failure modes which need to be diagnosed by the heuristic expert knowledge. The heuristic expert knowledge is implemented using a real-time expert system tool called G2 by Gensym Corp. Finally, the distributed diagnostic system requires another level of intelligence to oversee the fault mode reports generated by component fault detectors. The decision making at this level can best be done using a rule-based expert system. This level of expert knowledge is also implemented using G2.

Analysis of ProSEDS Test of Bare-Tether Collection

NASA Technical Reports Server (NTRS)

Sanmartin, J. R.; Lorenzini, E. C.; Estes, R. D.; Charro, M.; Cosmo, M. L.

2003-01-01

NASA's tether experiment ProSEDS will be placed in orbit on board a Delta-II rocket to test bare-tether electron collection, deorbiting of the rocket second stage, and the system dynamic stability. ProSEDS performance will vary because ambient conditions change along the orbit and tether-circuit bulk elements at the cathodic end follow the step-by-step sequence for the current cycles of operating modes (open-circuit, shunt and resistor modes for primary cycles; shunt and battery modes for secondary cycles). In this work we discuss expected ProSEDS values of the ratio L,/L*, which jointly with cathodic bulk elements determines bias and current tether profiles; L, is tether length, and L* (changing with tether temperature and ionospheric plasma density and magnetic field) is a characteristic length gauging ohmic versus baretether collection impedances. We discuss how to test bare-tether electron collection during primary cycles, using probe measurements of plasma density, measurements of cathodic current in resistor and shunt modes, and an estimate of tether temperature based on ProSEDS orbital position at the particular cycle concerned. We discuss how a temperature misestimate might occasionally affect the test of bare-tether collection, and how introducing the battery mode in some primary cycles, for an additional current measurement, could obviate the need of a temperature estimate. We also show how to test bare-tether collection by estimating orbit-decay rate from measurements of cathodic current for the shunt and battery modes of secondary cycles.

Wall Pressure Unsteadiness and Side Loads in Overexpanded Rocket Nozzles

NASA Technical Reports Server (NTRS)

Baars, Woutijn J.; Tinney, Charles E.; Ruf, Joseph H.; Brown, Andrew M.; McDaniels, David M.

2012-01-01

Surveys of both the static and dynamic wall pressure signatures on the interior surface of a sub-scale, cold-flow and thrust optimized parabolic nozzle are conducted during fixed nozzle pressure ratios corresponding to FSS and RSS states. The motive is to develop a better understanding for the sources of off-axis loads during the transient start-up of overexpanded rocket nozzles. During FSS state, pressure spectra reveal frequency content resembling SWTBLI. Presumably, when the internal flow is in RSS state, separation bubbles are trapped by shocks and expansion waves; interactions between the separated flow regions and the waves produce asymmetric pressure distributions. An analysis of the azimuthal modes reveals how the breathing mode encompasses most of the resolved energy and that the side load inducing mode is coherent with the response moment measured by strain gauges mounted upstream of the nozzle on a flexible tube. Finally, the unsteady pressure is locally more energetic during RSS, albeit direct measurements of the response moments indicate higher side load activity when in FSS state. It is postulated that these discrepancies are attributed to cancellation effects between annular separation bubbles.

The Alfred Nobel rocket camera. An early aerial photography attempt

NASA Astrophysics Data System (ADS)

Ingemar Skoog, A.

2010-02-01

Alfred Nobel (1833-1896), mainly known for his invention of dynamite and the creation of the Nobel Prices, was an engineer and inventor active in many fields of science and engineering, e.g. chemistry, medicine, mechanics, metallurgy, optics, armoury and rocketry. Amongst his inventions in rocketry was the smokeless solid propellant ballistite (i.e. cordite) patented for the first time in 1887. As a very wealthy person he actively supported many Swedish inventors in their work. One of them was W.T. Unge, who was devoted to the development of rockets and their applications. Nobel and Unge had several rocket patents together and also jointly worked on various rocket applications. In mid-1896 Nobel applied for patents in England and France for "An Improved Mode of Obtaining Photographic Maps and Earth or Ground Measurements" using a photographic camera carried by a "…balloon, rocket or missile…". During the remaining of 1896 the mechanical design of the camera mechanism was pursued and cameras manufactured. In April 1897 (after the death of Alfred Nobel) the first aerial photos were taken by these cameras. These photos might be the first documented aerial photos taken by a rocket borne camera. Cameras and photos from 1897 have been preserved. Nobel did not only develop the rocket borne camera but also proposed methods on how to use the photographs taken for ground measurements and preparing maps.

Airbreathing combined cycle engine systems

NASA Technical Reports Server (NTRS)

Rohde, John

1992-01-01

The Air Force and NASA share a common interest in developing advanced propulsion systems for commercial and military aerospace vehicles which require efficient acceleration and cruise operation in the Mach 4 to 6 flight regime. The principle engine of interest is the turboramjet; however, other combined cycles such as the turboscramjet, air turborocket, supercharged ejector ramjet, ejector ramjet, and air liquefaction based propulsion are also of interest. Over the past months careful planning and program implementation have resulted in a number of development efforts that will lead to a broad technology base for those combined cycle propulsion systems. Individual development programs are underway in thermal management, controls materials, endothermic hydrocarbon fuels, air intake systems, nozzle exhaust systems, gas turbines and ramjet ramburners.

Simplified, inverse, ejector design tool

NASA Technical Reports Server (NTRS)

Dechant, Lawrence J.

1993-01-01

A simple lumped parameter based inverse design tool has been developed which provides flow path geometry and entrainment estimates subject to operational, acoustic, and design constraints. These constraints are manifested through specification of primary mass flow rate or ejector thrust, fully-mixed exit velocity, and static pressure matching. Fundamentally, integral forms of the conservation equations coupled with the specified design constraints are combined to yield an easily invertible linear system in terms of the flow path cross-sectional areas. Entrainment is computed by back substitution. Initial comparison with experimental and analogous one-dimensional methods show good agreement. Thus, this simple inverse design code provides an analytically based, preliminary design tool with direct application to High Speed Civil Transport (HSCT) design studies.

P and W propulsion systems studies results/status

NASA Technical Reports Server (NTRS)

Smith, Martin G., Jr.; Champagne, George A.

1992-01-01

The topics covered include the following: Pratt and Whitney (P&W) propulsion systems studies - NASA funded efforts to date; P&W engine concepts; P&W combustor focus - rich burn quick quench (RBQQ) concept; mixer ejector nozzle concept - large flow entrainment reduces jet noise; technology impact on NO(x) emissions - mature RBQQ combustor reduces NO(x) up to 85 percent; technology impact on sideline noise characteristics of Mach 2.4 turbine bypass engines (TBE's) - 600 lb/sec airflow size; technology impact on takeoff gross weight (TOGW) - provides up to 12 percent TOGW reduction; HSCT quiet engine concepts; TBE inlet valve/ejector nozzle concept schematic; mixed flow turbofan study; and exhaust nozzle conceptual design.

Capture-ejector satellites

NASA Technical Reports Server (NTRS)

Macconochie, I. O.; Eldred, C. H.; Martin, J. A.

1983-01-01

A satellite in the form of a large rotating rim which can be used to boost spacecraft from low-Earth orbit to higher orbits is described. The rim rotates in the plane of its orbit such that the lower portion of the rim is traveling at suborbital velocity, while the upper portion is travelling at greater than orbital velocity. Ascending spacecraft or payloads arrive at the lowest portion of the rim at suborbital velocities, where the payloads are released on a trajectory for higher orbits; descending payloads employ the reverse procedure. Electric thrusters placed on the rim maintain rim rotational speed and altitude. From the standpoint of currently known materials, the capture-ejector concept may be useful for relatively small velocity increments.

Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

NASA Image and Video Library

1999-08-25

A close-up view of the front end of a Pegasus rocket booster being prepared by technicians at the Dryden Flight Research Center for flight tests with the X-43A "Hypersonic Experimental Vehicle," or "Hyper-X." The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude).

Three-dimensional finite element analysis of acoustic instability of solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Hackett, R. M.; Juruf, R. S.

1976-01-01

A three dimensional finite element solution of the acoustic vibration problem in a solid propellant rocket motor is presented. The solution yields the natural circular frequencies of vibration and the corresponding acoustic pressure mode shapes, considering the coupled response of the propellant grain to the acoustic oscillations occurring in the motor cavity. The near incompressibility of the solid propellant is taken into account in the formulation. A relatively simple example problem is solved in order to illustrate the applicability of the analysis and the developed computer code.

Nonlinear Combustion Instability Prediction

NASA Technical Reports Server (NTRS)

Flandro, Gary

2010-01-01

The liquid rocket engine stability prediction software (LCI) predicts combustion stability of systems using LOX-LH2 propellants. Both longitudinal and transverse mode stability characteristics are calculated. This software has the unique feature of being able to predict system limit amplitude.

Space Shuttle Redesigned Solid Rocket Motor nozzle natural frequency variations with burn time

NASA Technical Reports Server (NTRS)

Lui, C. Y.; Mason, D. R.

1991-01-01

The effects of erosion and thermal degradation on the Space Shuttle Redesigned Solid Rocket Motor (RSRM) nozzle's structural dynamic characteristics were analytically evaluated. Also considered was stiffening of the structure due to internal pressurization. A detailed NASTRAN finite element model of the nozzle was developed and used to evaluate the influence of these effects at several discrete times during motor burn. Methods were developed for treating erosion and thermal degradation, and a procedure was developed to account for internal pressure stiffening using differential stiffness matrix techniques. Results were verified using static firing test accelerometer data. Fast Fourier Transform and Maximum Entropy Method techniques were applied to the data to generate waterfall plots which track modal frequencies with burn time. Results indicate that the lower frequency nozzle 'vectoring' modes are only slightly affected by erosion, thermal effects and internal pressurization. The higher frequency shell modes of the nozzle are, however, significantly reduced.

Detailed Measurement of ORSC Main Chamber Injector Dynamics

NASA Astrophysics Data System (ADS)

Bedard, Michael J.

Improving fidelity in simulation of combustion dynamics in rocket combustors requires an increase in experimental measurement fidelity for validation. In a model rocket combustor, a chemiluminescence based spectroscopy technique was used to capture flame light emissions for direct comparison to a computational simulation of the production of chemiluminescent species. The comparison indicated that high fidelity models of rocket combustors can predict spatio-temporal distribution of chemiluminescent species with trend-wise accuracy. The comparison also indicated the limited ability of OH* and CH* emission to indicate flame heat release. Based on initial spectroscopy experiments, a photomultiplier based chemiluminescence sensor was designed to increase the temporal resolution of flame emission measurements. To apply developed methodologies, an experiment was designed to investigate the flow and combustion dynamics associated with main chamber injector elements typical of the RD-170 rocket engine. A unique feature of the RD-170 injector element is the beveled expansion between the injector recess and combustion chamber. To investigate effects of this geometry, a scaling methodology was applied to increase the physical scale of a single injector element while maintaining traceability to the RD-170 design. Two injector configurations were tested, one including a beveled injector face and the other a flat injector face. This design enabled improved spatial resolution of pressure and light emission measurements densely arranged in the injector recess and near-injector region of the chamber. Experimental boundary conditions were designed to closely replicate boundary conditions in simulations. Experimental results showed that the beveled injector face had a damping effect on pressure fluctuations occurring near the longitudinal resonant acoustic modes of the chamber, implying a mechanism for improved overall combustion stability. Near the injector, the beveled geometry resulted in more acoustic energy into higher frequency modes, while the flat-face geometry excited modes closer to the fundamental longitudinal mode frequency and its harmonics. Multi-scale analysis techniques were used to investigate intermittency and the range of physical scales present in measured signals. Flame light emission measurements confirmed the presence of flame holding in the injector recess in both configurations. Analysis of dynamics in light emission signals showed flame response at the chamber acoustic resonance frequency in addition to non-acoustic modes associated with mixing shear layer dynamics in the injector recess. The first known benchmark quality data sets of such injector dynamics were recorded in each configuration to enable pressure-based validation of high fidelity models of gas-centered swirl coaxial injectors. This work presents a critical contribution to development of validated combustion dynamics predictive tools and to the understanding of gas-centered swirl coaxial injector elements.

Technology for Transient Simulation of Vibration during Combustion Process in Rocket Thruster

NASA Astrophysics Data System (ADS)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2018-01-01

The article describes the technology for simulation of transient combustion processes in the rocket thruster for determination of vibration frequency occurs during combustion. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. The way to generate the Flamelet library with CFX-RIF was described. A technique for modeling transient combustion processes in the rocket thruster was proposed based on the Flamelet library. A cyclic irregularity of the temperature field like vortex core precession was detected in the chamber. Frequency of flame precession was obtained with the proposed simulation technique.

PARC Analysis of the NASA/GE 2D NRA Mixer/Ejector Nozzle

NASA Technical Reports Server (NTRS)

DeBonis, J. R.

1999-01-01

Interest in developing a new generation supersonic transport has increased in the past several years. Current projections indicate this aircraft would cruise at approximately Mach 2.4, have a range of 5000 nautical miles and carry at least 250 passengers. A large market for such an aircraft will exist in the next century due to a predicted doubling of the demand for long range air transportation by the end of the century and the growing influence of the Pacific Rim nations. Such a proposed aircraft could more than halve the flying time from Los Angeles to Tokyo. However, before a new economically feasible supersonic transport can be built, many key technologies must be developed. Among these technologies is noise suppression. Propulsion systems for a supersonic transport using current technology would exceed acceptable noise levels. All new aircraft must satisfy FAR 36 Stage III noise regulations. The largest area of concern is the noise generated during takeoff. A concerted effort under NASA's High Speed Research (HSR) program has begun to address the problem of noise suppression. One of the most promising concepts being studied in the area of noise suppression is the mixer/ejector nozzle. This study analyzes a typical noise suppressing mixer ejector nozzle at take off conditions, using a Full Navier-Stokes (FNS) computational fluid dynamics (CFD) code.

Parametric Study of a Mixer/Ejector Nozzle with Mixing Enhancement Devices

NASA Technical Reports Server (NTRS)

DalBello, T.; Steffen, C. J., Jr.

2001-01-01

A numerical study employing a simplified model of the High Speed Civil Transport mixer/ejector nozzle has been conducted to investigate the effect of tabs (vortex generators) on the mixing process. More complete mixing of the primary and secondary flows within the confined ejector lowers peak exit velocity resulting in reduced jet noise. Tabs were modeled as vortex pairs and inserted into the computational model. The location, size, and number of tabs were varied and its effect on the mixing process is presented here both quantitatively and qualitatively. A baseline case (no tabs) along with six other cases involving two different vortex strengths at three different orientations have been computed and analyzed. The case with the highest vorticity (six vortices representing large tabs) gives the best mixing. It is shown that the influence of the vorticity acts primarily in the forward or middle portions of the duct, significantly alters the flow structure, and promotes some mixing in the lateral direction. Unmixed pockets were found at the top and bottom of the lobe, and more clever placement of tabs improved mixing in the vertical direction. The technique of replacing tabs with vortices shows promise as an efficient tool for quickly optimizing tab placement in lobed mixers.

Resonant Interaction of a Linear Array of Supersonic Rectangular Jets: an Experimental Study

NASA Technical Reports Server (NTRS)

Raman, Ganesh; Taghavi, Ray

1994-01-01

This paper examines a supersonic multi jet interaction problem that we believe is likely to be important for mixing enhancement and noise reduction in supersonic mixer-ejector nozzles. We demonstrate that it is possible to synchronize the screech instability of four rectangular jets by precisely adjusting the inter jet spacing. Our experimental data agrees with a theory that assumes that the phase-locking of adjacent jets occurs through a coupling at the jet lip. Although the synchronization does not change the frequency of the screech tone, its amplitude is augmented by 10 dB. The synchronized multi jets exhibit higher spreading than the unsynchronized jets, with the single jet spreading the least. We compare the nearfield noise of the four jets with synchronized screech to the noise of the sum of four jets operated individually. Our noise measurements reveal that the more rapid mixing of the synchronized multi jets causes the peak jet noise source to move up stream and to radiate noise at larger angles to the flow direction. Based on our results, we believe that screech synchronization is advantageous for noise reduction internal to a mixer-ejector nozzle, since the noise can now be suppressed by a shorter acoustically lined ejector.

Static and wind tunnel near-field/far-field jet noise measurements from model scale single-flow base line and suppressor nozzles. Summary report. [conducted in the Boeing large anechoic test chamber and the NASA-Ames 40by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Jaeck, C. L.

1977-01-01

A test program was conducted in the Boeing large anechoic test chamber and the NASA-Ames 40- by 80-foot wind tunnel to study the near- and far-field jet noise characteristics of six baseline and suppressor nozzles. Static and wind-on noise source locations were determined. A technique for extrapolating near field jet noise measurements into the far field was established. It was determined if flight effects measured in the near field are the same as those in the far field. The flight effects on the jet noise levels of the baseline and suppressor nozzles were determined. Test models included a 15.24-cm round convergent nozzle, an annular nozzle with and without ejector, a 20-lobe nozzle with and without ejector, and a 57-tube nozzle with lined ejector. The static free-field test in the anechoic chamber covered nozzle pressure ratios from 1.44 to 2.25 and jet velocities from 412 to 594 m/s at a total temperature of 844 K. The wind tunnel flight effects test repeated these nozzle test conditions with ambient velocities of 0 to 92 m/s.

Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

NASA Technical Reports Server (NTRS)

Nielsen, D.; Townsend, J.; Kappus, K.; Driskill, T.; Torres, I.; Parks, R.

2005-01-01

The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.