Pulsations Induced by Vibrations in Aircraft Engine Two-Stage Pump

NASA Astrophysics Data System (ADS)

Gafurov, S. A.; Salmina, V. A.; Handroos, H.

2018-01-01

This paper describes a phenomenon of induced pressure pulsations inside a two-stage aircraft engine pump. A considered pumps consists of a screw-centrifugal and gear stages. The paper describes the cause of two-stage pump elements loading. A number of hypothesis of pressure pulsations generation inside a pump were considered. The main focus in this consideration is made on phenomena that are not related to pump mode of operation. Provided analysis has shown that pump vibrations as well as pump elements self-oscillations are the main causes that lead to trailing vortices generation. Analysis was conducted by means FEM and CFD simulations as well by means of experimental investigations to obtain natural frequencies and flow structure inside a screw-centrifugal stage. To perform accurate simulations adequate boundary conditions were considered. Cavitation and turbulence phenomena have been also taken into account. Obtained results have shown generated trailing vortices lead to high-frequency loading of the impeller of screw-centrifugal stage and can be a cause of the bearing damage.

Beta 2: A near term, fully reusable, horizontal takeoff and landing two-stage-to-orbit launch vehicle concept

NASA Technical Reports Server (NTRS)

Burkardt, Leo A.

1992-01-01

A recent study has confirmed the feasibility of a near term, fully reusable, horizontal takeoff and landing two-stage-to-orbit (TSTO) launch vehicle concept. The vehicle stages at Mach 6.5. The first stage is powered by a turboramjet propulsion system with the turbojets being fueled by JP and the ramjet by LH2. The second stage is powered by a space shuttle main engine (SSME) rocket engine. For about the same gross weight as growth versions of the 747, the vehicle can place 10,000 lbm. in low polar orbit or 16,000 lbm. to Space Station Freedom.

The design of two stage to orbit vehicles

NASA Astrophysics Data System (ADS)

Gregorek, G. M.; Ramsay, T. N.

1991-09-01

Two designs are presented for a two-stage-to-orbit vehicle to complement an existing heavy lift vehicle. The payload is 10,000 lbs and 27 ft long by 10 ft in diameter for design purposes and must be carried to a low earth orbit by an air-breathing carrier configuration that can take off horizontally within 15,000 ft. Two designs are presented: a delta wing/body carrier in which the fuselage contains the orbiter; and a cranked-delta wing/body carrier in which the orbiter is carried piggy back. The engines for both carriers are turbofanramjets powered with liquid hydrogen, and the orbiters employ either a Space Shuttle Main Engine or a half-scale version with additional scramjet engines. The orbiter based on a full-scale Space Shuttle Main Engine is found to have a significantly higher takeoff weight which results in a higher total takeoff weight.

Prediction of the Main Engine Power of a New Container Ship at the Preliminary Design Stage

NASA Astrophysics Data System (ADS)

Cepowski, Tomasz

2017-06-01

The paper presents mathematical relationships that allow us to forecast the estimated main engine power of new container ships, based on data concerning vessels built in 2005-2015. The presented approximations allow us to estimate the engine power based on the length between perpendiculars and the number of containers the ship will carry. The approximations were developed using simple linear regression and multivariate linear regression analysis. The presented relations have practical application for estimation of container ship engine power needed in preliminary parametric design of the ship. It follows from the above that the use of multiple linear regression to predict the main engine power of a container ship brings more accurate solutions than simple linear regression.

Evaluation of a staged fuel combustor for turboprop engines

NASA Technical Reports Server (NTRS)

Verdouw, A. J.

1976-01-01

Proposed EPA emission regulations require emission reduction by 1979 for various gas turbine engine classes. Extensive combustion technology advancements are required to meet the proposed regulations. The T56 turboprop engine requires CO, UHC, and smoke reduction. A staged fuel combustor design was tested on a combustion rig to evaluate emission reduction potential in turboprop engines from fuel zoning. The can-type combustor has separately fueled-pilot and main combustion zones in series. The main zone fueling system was arranged for potential incorporation into the T56 with minor or no modifications to the basic engine. Three combustor variable geometry systems were incorporated to evaluate various airflow distributions. Emission results with fixed geometry operation met all proposed EPA regulations over the EPA LTO cycle. CO reduction was 82 percent, UHC reduction was 96 percent, and smoke reduction was 84 percent. NOx increased 14 percent over the LTO cycle. At high power, NOx reduction was 40 to 55 percent. This NOx reduction has potential application to stationary gas turbine powerplants which have different EPA regulations.

NASA Collaborative Design Processes

NASA Technical Reports Server (NTRS)

Jones, Davey

2017-01-01

This is Block 1, the first evolution of the world's most powerful and versatile rocket, the Space Launch System, built to return humans to the area around the moon. Eventually, larger and even more powerful and capable configurations will take astronauts and cargo to Mars. On the sides of the rocket are the twin solid rocket boosters that provide more than 75 percent during liftoff and burn for about two minutes, after which they are jettisoned, lightening the load for the rest of the space flight. Four RS-25 main engines provide thrust for the first stage of the rocket. These are the world's most reliable rocket engines. The core stage is the main body of the rocket and houses the fuel for the RS-25 engines, liquid hydrogen and liquid oxygen, and the avionics, or "brain" of the rocket. The core stage is all new and being manufactured at NASA's "rocket factory," Michoud Assembly Facility near New Orleans. The Launch Vehicle Stage Adapter, or LVSA, connects the core stage to the Interim Cryogenic Propulsion Stage. The Interim Cryogenic Propulsion Stage, or ICPS, uses one RL-10 rocket engine and will propel the Orion spacecraft on its deep-space journey after first-stage separation. Finally, the Orion human-rated spacecraft sits atop the massive Saturn V-sized launch vehicle. Managed out of Johnson Space Center in Houston, Orion is the first spacecraft in history capable of taking humans to multiple destinations within deep space. 2) Each element of the SLS utilizes collaborative design processes to achieve the incredible goal of sending human into deep space. Early phases are focused on feasibility and requirements development. Later phases are focused on detailed design, testing, and operations. There are 4 basic phases typically found in each phase of development.

Hot-Fire Test Results of Liquid Oxygen/RP-2 Multi-Element Oxidizer-Rich Preburners

NASA Technical Reports Server (NTRS)

Protz, C. S.; Garcia, C. P.; Casiano, M. J.; Parton, J. A.; Hulka, J. R.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. To supply the oxidizer-rich combustion products to the main injector of the integrated test article, existing subscale preburner injectors from a previous NASA-funded oxidizer-rich staged combustion engine development program were utilized. For the integrated test article, existing and newly designed and fabricated inter-connecting hot gas duct hardware were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. However, before one of the preburners was used in the integrated test article, it was first hot-fire tested at length to prove it could provide the hot exhaust gas mean temperature, thermal uniformity and combustion stability necessary to perform in the integrated test article experiment. This paper presents results from hot-fire testing of several preburner injectors in a representative combustion chamber with a sonic throat. Hydraulic, combustion performance, exhaust gas thermal uniformity, and combustion stability data are presented. Results from combustion stability modeling of these test results are described in a companion paper at this JANNAF conference, while hot-fire test results of the preburner injector in the integrated test article are described in another companion paper.

Full Hybrid: Stopped

Science.gov Websites

than batteries used to start conventional vehicles. Main stage: See through car with battery, engine electric motor or stored in the battery. It is also used to start the gasoline engine instantly when needed used to start conventional vehicles

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

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.

J-2X, The Engine of the Future

NASA Technical Reports Server (NTRS)

Smith, Gail

2009-01-01

My project was two-fold, with both parts involving the J-2X Upper Stage engine (which will be used on both the Ares I and V). Mainly, I am responsible for using a program called Iris to create visual represen tations of the rocket engine's telemetry data. Also, my project includes the application of my newly acquired Pro Engineer skills in develo ping a 3D model of the engine's nozzle.

Photographic copy of drawing by Modjeski and Masters, Engineers of ...

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

Photographic copy of drawing by Modjeski and Masters, Engineers of the proposed Huey P. Long Bridge Widening. Original drawing located in the office of Modjeski and Masters, Consulting Engineers at 1055 St. Charles Avenue, New Orleans, LA. 70130. JUNE 30, 2004 DRAWING OF THE PROPOSED HUEY P. LONG BRIDGE WIDENING, U.S. 90, MAIN BRIDGE SUPERSTRUCTURE, SHOWING SEQUENCE OF CONSTRUCTION – 16, STAGE 5 (PLACEMENT OF NEW DECK AND FLOOR SYSTEM) AND STAGE 6. - Huey P. Long Bridge, Spanning Mississippi River approximately midway between nine & twelve mile points upstream from & west of New Orleans, Jefferson, Jefferson Parish, LA

Photographic copy of drawing by Modjeski and Masters, Engineers of ...

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

Photographic copy of drawing by Modjeski and Masters, Engineers of the proposed Huey P. Long Bridge Widening. Original drawing located in the office of Modjeski and Masters, Consulting Engineers at 1055 St. Charles Avenue, New Orleans, LA. 70130. JUNE 30, 2004 DRAWING OF THE PROPOSED HUEY P. LONG BRIDGE WIDENING, U.S. 90, MAIN BRIDGE SUPERSTRUCTURE, SHOWING SEQUENCE OF CONSTRUCTION – 14, STAGE 3 – PHASE 1 AND STAGE 3 – PHASE 2. - Huey P. Long Bridge, Spanning Mississippi River approximately midway between nine & twelve mile points upstream from & west of New Orleans, Jefferson, Jefferson Parish, LA

Photographic copy of drawing by Modjeski and Masters, Engineers of ...

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

Photographic copy of drawing by Modjeski and Masters, Engineers of the proposed Huey P. Long Bridge Widening. Original drawing located in the office of Modjeski and Masters, Consulting Engineers at 1055 St. Charles Avenue, New Orleans, LA. 70130. JUNE 30, 2004 DRAWING OF THE PROPOSED HUEY P. LONG BRIDGE WIDENING, U.S. 90, MAIN BRIDGE SUPERSTRUCTURE, SHOWING SEQUENCE OF CONSTRUCTION – 15, STAGE 4 AND STAGE 5 (DECK AND FLOOR SYSTEM REMOVAL). - Huey P. Long Bridge, Spanning Mississippi River approximately midway between nine & twelve mile points upstream from & west of New Orleans, Jefferson, Jefferson Parish, LA

Thermoacoustic refrigerators and engines comprising cascading stirling thermodynamic units

DOEpatents

Backhaus, Scott; Swift, Greg

2013-06-25

The present invention includes a thermoacoustic assembly and method for improved efficiency. The assembly has a first stage Stirling thermal unit comprising a main ambient heat exchanger, a regenerator and at least one additional heat exchanger. The first stage Stirling thermal unit is serially coupled to a first end of a quarter wavelength long coupling tube. A second stage Stirling thermal unit comprising a main ambient heat exchanger, a regenerator, and at least one additional heat exchanger, is serially coupled to a second end of the quarter wavelength long coupling tube.

Chinese modify CZ-2/3 rocket boosters, focus on commercial launch market

NASA Astrophysics Data System (ADS)

Covault, C.

1985-07-01

A program underway in the People's Republic of China to modify the Titan-class CZ-2/3 satellite-launch and ICBM boosters is described on the basis of a recent visit to the manufacturing plant in Shanghai. The present two-stage CZ-2 and three-stage CZ-3 can place 5000 lbs in LEO or 3080 lbs in GEO, respectively, and are produced on a custom basis with a delivery time of about 2 yrs. Modifications introduced include 4 x 6-ft fins and a pogo-suppression system for the four-engine first stage and a steel support band for the combustion chamber of the 80-ton-thrust second-stage main engine.

Dual-Fuel Propulsion in Single-Stage Advanced Manned Launch System Vehicle

NASA Technical Reports Server (NTRS)

Lepsch, Roger A., Jr.; Stanley, Douglas O.; Unal, Resit

1995-01-01

As part of the United States Advanced Manned Launch System study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined. Several dual-fuel propulsion concepts were investigated. These include: a separate-engine concept combining Russian RD-170 kerosene-fueled engines with space shuttle main engine-derivative engines: the kerosene- and hydrogen-fueled Russian RD-701 engine; and a dual-fuel, dual-expander engine. Analysis to determine vehicle weight and size characteristics was performed using conceptual-level design techniques. A response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight. The tools and methods employed in the analysis process are also summarized. In comparison with a reference hydrogen- fueled single-stage vehicle, results showed that the dual-fuel vehicles were from 10 to 30% lower in empty weight for the same payload capability, with the dual-expander engine types showing the greatest potential.

Development of a CFRP Engine Thrust Frame for the Next Generation Launchers

NASA Astrophysics Data System (ADS)

Fatemi, Javad; van der Bas, Finn; Cruijssen, Henk

2012-07-01

This paper addresses the activities related to the development of technologies for a composite Engine Thrust Frame (ETF) for the next generation launchers. In particular, the design and analyses of a full Carbon Fibre Reinforced Plastic (CFRP) engine thrust frame are presented in more detail. The ETF concept is composed of three main parts, i.e. an aluminium top-ring which connects the ETF to the upper-stage tank, a CFRP cone, and a CFRP cone-cap which connects the Vinci engine to the ETF. The main challenging requirements for development of a CFRP ETF are recalled. The ETF concept and its mechanical performances are assessed.

Last SSME test on A-1

NASA Image and Video Library

2006-09-29

The Stennis Space Center conducted the final space shuttle main engine test on its A-1 Test Stand Friday. The A-1 Test Stand was the site of the first test on a shuttle main engine in 1975. Stennis will continue testing shuttle main engines on its A-2 Test Stand through the end of the Space Shuttle Program in 2010. The A-1 stand begins a new chapter in its operational history in October. It will be temporarily decommissioned to convert it for testing the J-2X engine, which will power the upper stage of NASA's new crew launch vehicle, the Ares I. Although this ends the stand's work on the Space Shuttle Program, it will soon be used for the rocket that will carry America's next generation human spacecraft, Orion.

Emission Characteristics of A P and W Axially Staged Sector Combustor

NASA Technical Reports Server (NTRS)

He, Zhuohui J.; Wey, Changlie; Chang, Clarence T.; Lee, Chi Ming; Surgenor, Angela D.; Kopp-Vaughan, Kristin; Cheung, Albert

2016-01-01

Emission characteristics of a three-cup P and W Axially Controlled Stoichiometry (ACS) sector combustor are reported in this article. Multiple injection points and fuel staging strategies are used in this combustor design. Pilot-stage injectors are located on the front dome plate of the combustor, and main-stage injectors are positioned on the top and bottom of the combustor liners downstream. Low power configuration uses only pilot-stage injectors. Main-stage injectors are added to high power configuration to help distribute fuel more evenly and achieve overall lean burn yielding very low NOx emissions. Combustion efficiencies at four ICAO LTO conditions were all above 99%. Three EINOx emissions correlation equations were developed based on the experimental data to describe the NOx emission trends of this combustor concept. For the 7% and 30% engine power conditions, NOx emissions are obtained with the low power configuration, and the EINOx values are 6.16 and 6.81. The high power configuration was used to assess 85% and 100% engine power NOx emissions, with measured EINOx values of 4.58 and 7.45, respectively. The overall landing-takeoff cycle NOx emissions are about 12% relative to ICAO CAEP/6 level.

Celebrating 50 Years of Testing

NASA Image and Video Library

2016-04-19

What better way to mark 50 years of rocket engine testing than with a rocket engine test? Stennis Space Center employees enjoyed a chance to view an RS-68 engine test at the B-1 Test Stand on April 19, almost 50 years to the day that the first test was conducted at the south Mississippi site in 1966. The test viewing was part of a weeklong celebration of the 50th year of rocket engine testing at Stennis. The first test at the site occurred April 23, 1966, with a 15-second firing of a Saturn V second stage prototype (S-II-C) on the A-2 Test Stand. The center subsequently tested Apollo rocket stages that carried humans to the moon and every main engine used to power 135 space shuttle missions. It currently tests engines for NASA’s new Space Launch System vehicle.

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

Not Available

The model is designed to enable decision makers to compare the economics of geothermal projects with the economics of alternative energy systems at an early stage in the decision process. The geothermal engineering and economic feasibility computer model (GEEF) is written in FORTRAN IV language and can be run on a mainframe or a mini-computer system. An abbreviated version of the model is being developed for usage in conjunction with a programmable desk calculator. The GEEF model has two main segments, namely (i) the engineering design/cost segment and (ii) the economic analysis segment. In the engineering segment, the model determinesmore » the numbers of production and injection wells, heat exchanger design, operating parameters for the system, requirement of supplementary system (to augment the working fluid temperature if the resource temperature is not sufficiently high), and the fluid flow rates. The model can handle single stage systems as well as two stage cascaded systems in which the second stage may involve a space heating application after a process heat application in the first stage.« less

Stop/Start: Driving

Science.gov Websites

/generator visible. The car is moving. There are purple arrows flowing from the gasoline engine to the electric starter/generator. There are red arrows flowing from the gasoline engine to the front wheels . There are blue arrows flowing from the electric starter/generator to the battery. Main stage: See

Phase 1 Development Testing of the Advanced Manufacturing Demonstrator Engine

NASA Technical Reports Server (NTRS)

Case, Nicholas L.; Eddleman, David E.; Calvert, Marty R.; Bullard, David B.; Martin, Michael A.; Wall, Thomas R.

2016-01-01

The Additive Manufacturing Development Breadboard Engine (BBE) is a pressure-fed liquid oxygen/pump-fed liquid hydrogen (LOX/LH2) expander cycle engine that was built and operated by NASA at Marshall Space Flight Center's East Test Area. The breadboard engine was conceived as a technology demonstrator for the additive manufacturing technologies for an advanced upper stage prototype engine. The components tested on the breadboard engine included an ablative chamber, injector, main fuel valve, turbine bypass valve, a main oxidizer valve, a mixer and the fuel turbopump. All parts minus the ablative chamber were additively manufactured. The BBE was successfully hot fire tested seven times. Data collected from the test series will be used for follow on demonstration tests with a liquid oxygen turbopump and a regeneratively cooled chamber and nozzle.

Tripropellant engine study

NASA Technical Reports Server (NTRS)

Wheeler, D. B.; Kirby, F. M.

1978-01-01

The potential for converting the space shuttle main engine (SSME) to a dual-fuel, dual-mode engine using LOX/hydrocarbon propellants in mode 1 and LOX/H2 in mode 2 was examined. Various engine system concepts were formulated that included staged combustion and gas generator turbine power cycles, and LOX/RP-1, LOX/CH4, and LOX/C3H8 mode 1 propellants. Both oxidizer and fuel regenerative cooling were considered. All of the SSME major components were examined to determine their adaptability to the candidate dual-fuel engines.

Shuttle Propulsion Overview - The Design Challenges

NASA Technical Reports Server (NTRS)

Owen, James W.

2011-01-01

The major elements of the Space Shuttle Main Propulsion System include two reusable solid rocket motors integrated into recoverable solid rocket boosters, an expendable external fuel and oxidizer tank, and three reusable Space Shuttle Main Engines. Both the solid rocket motors and space shuttle main engines ignite prior to liftoff, with the solid rocket boosters separating about two minutes into flight. The external tank separates, about eight and a half minutes into the flight, after main engine shutdown and is safely expended in the ocean. The SSME's, integrated into the Space Shuttle Orbiter aft structure, are reused after post landing inspections. The configuration is called a stage and a half as all the propulsion elements are active during the boost phase, with only the SSME s continuing operation to achieve orbital velocity. Design and performance challenges were numerous, beginning with development work in the 1970's. The solid rocket motors were large, and this technology had never been used for human space flight. The SSME s were both reusable and very high performance staged combustion cycle engines, also unique to the Space Shuttle. The multi body side mount configuration was unique and posed numerous integration and interface challenges across the elements. Operation of the system was complex and time consuming. This paper describes the design challenges and key areas where the design evolved during the program.

Thermal finite-element analysis of space shuttle main engine turbine blade

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali; Tong, Michael T.; Kaufman, Albert

1987-01-01

Finite-element, transient heat transfer analyses were performed for the first-stage blades of the space shuttle main engine (SSME) high-pressure fuel turbopump. The analyses were based on test engine data provided by Rocketdyne. Heat transfer coefficients were predicted by performing a boundary-layer analysis at steady-state conditions with the STAN5 boundary-layer code. Two different peak-temperature overshoots were evaluated for the startup transient. Cutoff transient conditions were also analyzed. A reduced gas temperature profile based on actual thermocouple data was also considered. Transient heat transfer analyses were conducted with the MARC finite-element computer code.

Axially staged combustion system for a gas turbine engine

DOEpatents

Bland, Robert J [Oviedo, FL

2009-12-15

An axially staged combustion system is provided for a gas turbine engine comprising a main body structure having a plurality of first and second injectors. First structure provides fuel to at least one of the first injectors. The fuel provided to the one first injector is adapted to mix with air and ignite to produce a flame such that the flame associated with the one first injector defines a flame front having an average length when measured from a reference surface of the main body structure. Each of the second injectors comprising a section extending from the reference surface of the main body structure through the flame front and having a length greater than the average length of the flame front. Second structure provides fuel to at least one of the second injectors. The fuel passes through the one second injector and exits the one second injector at a location axially spaced from the flame front.

Overview of the Main Propulsion System for the NASA Ares I Upper Stage

NASA Technical Reports Server (NTRS)

Quinn, Jason E.; Swanson, Luke A.

2009-01-01

A functional overview of the Main Propulsion System (MPS) of the NASA Ares I Upper Stage is provided. In addition to a simple overview of the key MPS functions and design philosophies, major lessons learned are discussed. The intent is to provide a technical overview with enough detail to allow engineers outside of the MPS Integrated Product Team (IPT) to develop a rough understanding of MPS operations, components, design philosophy, and lessons learned.

Advanced Transportation System Studies. Technical Area 3: Alternate Propulsion Subsystem Concepts. Volume 1; Executive Summary

NASA Technical Reports Server (NTRS)

Levack, Daniel J. H.

2000-01-01

The Alternate Propulsion Subsystem Concepts contract had seven tasks defined that are reported under this contract deliverable. The tasks were: FAA Restart Study, J-2S Restart Study, Propulsion Database Development. SSME Upper Stage Use. CERs for Liquid Propellant Rocket Engines. Advanced Low Cost Engines, and Tripropellant Comparison Study. The two restart studies, F-1A and J-2S, generated program plans for restarting production of each engine. Special emphasis was placed on determining changes to individual parts due to obsolete materials, changes in OSHA and environmental concerns, new processes available, and any configuration changes to the engines. The Propulsion Database Development task developed a database structure and format which is easy to use and modify while also being comprehensive in the level of detail available. The database structure included extensive engine information and allows for parametric data generation for conceptual engine concepts. The SSME Upper Stage Use task examined the changes needed or desirable to use the SSME as an upper stage engine both in a second stage and in a translunar injection stage. The CERs for Liquid Engines task developed qualitative parametric cost estimating relationships at the engine and major subassembly level for estimating development and production costs of chemical propulsion liquid rocket engines. The Advanced Low Cost Engines task examined propulsion systems for SSTO applications including engine concept definition, mission analysis. trade studies. operating point selection, turbomachinery alternatives, life cycle cost, weight definition. and point design conceptual drawings and component design. The task concentrated on bipropellant engines, but also examined tripropellant engines. The Tripropellant Comparison Study task provided an unambiguous comparison among various tripropellant implementation approaches and cycle choices, and then compared them to similarly designed bipropellant engines in the SSTO mission This volume overviews each of the tasks giving its objectives, main results. and conclusions. More detailed Final Task Reports are available on each individual task.

Integrated Testing Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Taylor, James L.; Cockrell, Charles E.; Tuma, Margaret L.; Askins, Bruce R.; Bland, Jeff D.; Davis, Stephan R.; Patterson, Alan F.; Taylor, Terry L.; Robinson, Kimberly L.

2008-01-01

The Ares I crew launch vehicle is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew and cargo access to the International Space Station (ISS) and, together with the Ares V cargo launch vehicle, serves as a critical component of NASA's future human exploration of the Moon. During the preliminary design phase, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements - including the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine - will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the upper stage Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle ground vibration test (IVGVT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, validate the ability of the upper stage to manage cryogenic propellants to achieve upper stage engine start conditions, and a high-altitude demonstration of the launch abort system (LAS) following stage separation. The Orion 1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.

One Idea for a Next Generation Shuttle

NASA Technical Reports Server (NTRS)

MacConochie, Ian O.; Cerro, Jeffrey A.

2004-01-01

In this configuration, the current Shuttle External Tank serves as core structure for a fully reusable second stage. This stage is equipped with wings, vertical fin, landing gear, and thermal protection. The stage is geometrically identical to (but smaller than) a single stage that has been tested hyper-sonically, super-sonically, and sub-sonically in the NASA Langley Research Center wind tunnels. The three LOX/LH engines that currently serve as main propulsion for the Shuttle Orbiter, serve as main propulsion on the new stage. The new stage is unmanned but is equipped with the avionics needed for automatic maneuvering on orbit and for landing on a runway. Three rails are installed along the top surface of the vehicle for attachment of various payloads. Pay- loads might include third stages with satellites attached, personnel pods, propellants, or other items.

Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Protz, C. S.; Garcia, C. P.; Casiano, M. J.; Parton, J. A.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics.

Stop/Start: Overview

Science.gov Websites

/Start technology conserves energy by shutting off the gasoline engine when the vehicle is at rest, such as at a traffic light, and automatically re-starting it when the driver pushes the gas pedal to go engine when pulling out from a stop and generating electricity which is stored in the battery. Main stage

Building information modelling review with potential applications in tunnel engineering of China.

PubMed

Zhou, Weihong; Qin, Haiyang; Qiu, Junling; Fan, Haobo; Lai, Jinxing; Wang, Ke; Wang, Lixin

2017-08-01

Building information modelling (BIM) can be applied to tunnel engineering to address a number of problems, including complex structure, extensive design, long construction cycle and increased security risks. To promote the development of tunnel engineering in China, this paper combines actual cases, including the Xingu mountain tunnel and the Shigu Mountain tunnel, to systematically analyse BIM applications in tunnel engineering in China. The results indicate that BIM technology in tunnel engineering is currently mainly applied during the design stage rather than during construction and operation stages. The application of BIM technology in tunnel engineering covers many problems, such as a lack of standards, incompatibility of different software, disorganized management, complex combination with GIS (Geographic Information System), low utilization rate and poor awareness. In this study, through summary of related research results and engineering cases, suggestions are introduced and an outlook for the BIM application in tunnel engineering in China is presented, which provides guidance for design optimization, construction standards and later operation maintenance.

Building information modelling review with potential applications in tunnel engineering of China

PubMed Central

Zhou, Weihong; Qin, Haiyang; Fan, Haobo; Lai, Jinxing; Wang, Ke; Wang, Lixin

2017-01-01

Building information modelling (BIM) can be applied to tunnel engineering to address a number of problems, including complex structure, extensive design, long construction cycle and increased security risks. To promote the development of tunnel engineering in China, this paper combines actual cases, including the Xingu mountain tunnel and the Shigu Mountain tunnel, to systematically analyse BIM applications in tunnel engineering in China. The results indicate that BIM technology in tunnel engineering is currently mainly applied during the design stage rather than during construction and operation stages. The application of BIM technology in tunnel engineering covers many problems, such as a lack of standards, incompatibility of different software, disorganized management, complex combination with GIS (Geographic Information System), low utilization rate and poor awareness. In this study, through summary of related research results and engineering cases, suggestions are introduced and an outlook for the BIM application in tunnel engineering in China is presented, which provides guidance for design optimization, construction standards and later operation maintenance. PMID:28878970

Building information modelling review with potential applications in tunnel engineering of China

NASA Astrophysics Data System (ADS)

Zhou, Weihong; Qin, Haiyang; Qiu, Junling; Fan, Haobo; Lai, Jinxing; Wang, Ke; Wang, Lixin

2017-08-01

Building information modelling (BIM) can be applied to tunnel engineering to address a number of problems, including complex structure, extensive design, long construction cycle and increased security risks. To promote the development of tunnel engineering in China, this paper combines actual cases, including the Xingu mountain tunnel and the Shigu Mountain tunnel, to systematically analyse BIM applications in tunnel engineering in China. The results indicate that BIM technology in tunnel engineering is currently mainly applied during the design stage rather than during construction and operation stages. The application of BIM technology in tunnel engineering covers many problems, such as a lack of standards, incompatibility of different software, disorganized management, complex combination with GIS (Geographic Information System), low utilization rate and poor awareness. In this study, through summary of related research results and engineering cases, suggestions are introduced and an outlook for the BIM application in tunnel engineering in China is presented, which provides guidance for design optimization, construction standards and later operation maintenance.

Ares I Integrated Test Approach

NASA Technical Reports Server (NTRS)

Taylor, Jim

2008-01-01

This slide presentation reviews the testing approach that NASA is developing for the Ares I launch vehicle. NASA is planning a complete series of development, qualification and verification tests. These include: (1) Upper stage engine sea-level and altitude testing (2) First stage development and qualification motors (3) Upper stage structural and thermal development and qualification test articles (4) Main Propulsion Test Article (MPTA) (5) Upper stage green run testing (6) Integrated Vehicle Ground Vibration Testing (IVGVT) and (7) Aerodynamic characterization testing.

Determining Parameters of Double-Wiebe Function for Simulation of Combustion Process in Overload Diesel Engine with Common Rail Fuel Feed System

NASA Astrophysics Data System (ADS)

Kamaltdinov, V. G.; Markov, V. A.; Lysov, I. O.

2018-03-01

To analyze the peculiarities of the combustion process in an overload diesel engine with the system of Common Rail type with one-stage injection, the indicator diagram was registered. The parameters of the combustion process simulated by the double-Wiebe function were calculated as satisfactorily reconstructing the law of burning rate variation. The main parameters of the operating cycle obtained through the indicator diagram processing and the double-Wiebe function calculation differed insignificantly. And the calculated curve of the cylinder pressure differed notably only in the end of the expansion stroke. To improve the performance of the diesel engine, a two-stage fuel injection was recommended.

Independent Orbiter Assessment (IOA): Analysis of the ascent thrust vector control actuator subsystem

NASA Technical Reports Server (NTRS)

Wilson, R. E.; Riccio, J. R.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Ascent Thrust Vector Control (ATVC) Actuator hardware are documented. The function of the Ascent Thrust Vector Control Actuators (ATVC) is to gimbal the main engines to provide for attitude and flight path control during ascent. During first stage flight, the SRB nozzles provide nearly all the steering. After SRB separation, the Orbiter is steered by gimbaling of its main engines. There are six electrohydraulic servoactuators, one pitch and one yaw for each of the three main engines. Each servoactuator is composed of four electrohydraulic servovalve assemblies, one second stage power spool valve assembly, one primary piston assembly and a switching valve. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Critical failures resulting in loss of ATVC were mainly due to loss of hydraulic fluid, fluid contamination and mechanical failures.

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/ Hydrogen Rocket Engine

DTIC Science & Technology

1993-10-12

achieved on the same test. Figure 5-1 is a photograph of the RS-44 ICE engine as it achieved mainstage 3 operation on test 87-017-006. Nominal and...engine transition to mainstage main chamber pressure, 3 Igniter operation satisfactory 87-017-004 1.09 Igniter operation / engine Redline cutoff due...ok, transition stage satisfactory, 3 maInstage operation achieved, all other objectives met. 3 I I I -56- 1 CR 194443 Figure 7-22 RS-44 ENGtNE

Design feasibility via ascent optimality for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Miele, A.; Mancuso, S.

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor. The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload. For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).

Thermal-structural analyses of Space Shuttle Main Engine (SSME) hot section components

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali; Thompson, Robert L.

1988-01-01

Three dimensional nonlinear finite element heat transfer and structural analyses were performed for the first stage high pressure fuel turbopump (HPFTP) blade of the space shuttle main engine (SSME). Directionally solidified (DS) MAR-M 246 and single crystal (SC) PWA-1480 material properties were used for the analyses. Analytical conditions were based on a typical test stand engine cycle. Blade temperature and stress strain histories were calculated by using the MARC finite element computer code. The structural response of an SSME turbine blade was assessed and a greater understanding of blade damage mechanisms, convective cooling effects, and thermal mechanical effects was gained.

Design considerations for a Space Shuttle Main Engine turbine blade made of single crystal material

NASA Technical Reports Server (NTRS)

Abdul-Aziz, A.; August, R.; Nagpal, V.

1993-01-01

Nonlinear finite-element structural analyses were performed on the first stage high-pressure fuel turbopump blade of the Space Shuttle Main Engine. The analyses examined the structural response and the dynamic characteristics at typical operating conditions. Single crystal material PWA-1480 was considered for the analyses. Structural response and the blade natural frequencies with respect to the crystal orientation were investigated. The analyses were conducted based on typical test stand engine cycle. Influence of combined thermal, aerodynamic, and centrifugal loadings was considered. Results obtained showed that the single crystal secondary orientation effects on the maximum principal stresses are not highly significant.

Staged direct injection diesel engine

DOEpatents

Baker, Quentin A.

1985-01-01

A diesel engine having staged injection for using lower cetane number fuels than No. 2 diesel fuel. The engine includes a main fuel injector and a pilot fuel injector. Pilot and main fuel may be the same fuel. The pilot injector injects from five to fifteen percent of the total fuel at timings from 20.degree. to 180.degree. BTDC depending upon the quantity of pilot fuel injected, the fuel cetane number and speed and load. The pilot fuel injector is directed toward the centerline of the diesel cylinder and at an angle toward the top of the piston, avoiding the walls of the cylinder. Stratification of the early injected pilot fuel is needed to reduce the fuel-air mixing rate, prevent loss of pilot fuel to quench zones, and keep the fuel-air mixture from becoming too fuel lean to become effective. In one embodiment, the pilot fuel injector includes a single hole for injection of the fuel and is directed at approximately 48.degree. below the head of the cylinder.

Space Shuttle guidance for multiple main engine failures during first stage

NASA Technical Reports Server (NTRS)

Sponaugle, Steven J.; Fernandes, Stanley T.

1987-01-01

This paper presents contingency abort guidance schemes recently developed for multiple Space Shuttle main engine failures during the first two minutes of flight (first stage). The ascent and entry guidance schemes greatly improve the possibility of the crew and/or the Orbiter surviving a first stage contingency abort. Both guidance schemes were required to meet certain structural and controllability constraints. In addition, the systems were designed with the flexibility to allow for seasonal variations in the atmosphere and wind. The ascent scheme guides the vehicle to a desirable, lofted state at solid rocket booster burnout while reducing the structural loads on the vehicle. After Orbiter separation from the solid rockets and the external tank, the entry scheme guides the Orbiter through one of two possible entries. If the proper altitude/range/velocity conditions have been met, a return-to-launch-site 'Split-S' maneuver may be attempted. Otherwise, a down-range abort to an equilibrium glide and subsequent crew bailout is performed.

Crew Launch Vehicle (CLV) Upper Stage Configuration Selection Process

NASA Technical Reports Server (NTRS)

Davis, Daniel J.; Coook, Jerry R.

2006-01-01

The Crew Launch Vehicle (CLV), a key component of NASA's blueprint for the next generation of spacecraft to take humans back to the moon, is being designed and built by engineers at NASA s Marshall Space Flight Center (MSFC). The vehicle s design is based on the results of NASA's 2005 Exploration Systems Architecture Study (ESAS), which called for development of a crew-launch system to reduce the gap between Shuttle retirement and Crew Exploration Vehicle (CEV) Initial Operating Capability, identification of key technologies required to enable and significantly enhance these reference exploration systems, and a reprioritization of near- and far-term technology investments. The Upper Stage Element (USE) of the CLV is a clean-sheet approach that is being designed and developed in-house, with element management at MSFC. The USE concept is a self-supporting cylindrical structure, approximately 115' long and 216" in diameter, consisting of the following subsystems: Primary Structures (LOX Tank, LH2 Tank, Intertank, Thrust Structure, Spacecraft Payload Adaptor, Interstage, Forward and Aft Skirts), Secondary Structures (Systems Tunnel), Avionics and Software, Main Propulsion System, Reaction Control System, Thrust Vector Control, Auxiliary Power Unit, and Hydraulic Systems. The ESAS originally recommended a CEV to be launched atop a four-segment Space Shuttle Main Engine (SSME) CLV, utilizing an RS-25 engine-powered upper stage. However, Agency decisions to utilize fewer CLV development steps to lunar missions, reduce the overall risk for the lunar program, and provide a more balanced engine production rate requirement prompted engineers to switch to a five-segment design with a single Saturn-derived J-2X engine. This approach provides for single upper stage engine development for the CLV and an Earth Departure Stage, single Reusable Solid Rocket Booster (RSRB) development for the CLV and a Cargo Launch Vehicle, and single core SSME development. While the RSRB design has changed since the CLV Project's inception, the USE design has remained essentially a clean-sheet approach. Although a clean-sheet upper stage design inherently carries more risk than a modified design, it does offer many advantages: a design for increased reliability; built-in extensibility to allow for commonality/growth without major redesign; and incorporation of state-of-the-art materials, hardware, and design, fabrication, and test techniques and processes to facilitate a potentially better, more reliable system. Because consideration was given in the ESAS to both clean-sheet and modified USE designs, this paper will highlight the advantages and disadvantages of both approaches and provide a detailed discussion of trades/selections made that led to the final upper stage configuration.

Propulsion at the Marshall Space Flight Center - A brief history

NASA Technical Reports Server (NTRS)

Jones, L. W.; Fisher, M. F.; Mccool, A. A.; Mccarty, J. P.

1991-01-01

The history of propulsion development at the NASA Marshall Space Flight Center is summarized, beginning with the development of the propulsion system for the Redstone missile. This course of propulsion development continues through the Jupiter IRBM, the Saturn family of launch vehicles and the engines that powered them, the Centaur upper stage and RL-10 engine, the Reactor In-Flight Test stage and the NERVA nuclear engine. The Space Shuttle Main Engine and Solid Rocket Boosters are covered, as are spacecraft propulsion systems, including the reaction control systems for the High Energy Astronomy Observatory and the Space Station. The paper includes a description of several technology efforts such as those in high pressure turbomachinery, aerospike engines, and the AS203 cyrogenic fluid management flight experiment. These and other propulsion projects are documented, and the scope of activities in support of these efforts at Marshall delineated.

Multiscale/Multifunctional Probabilistic Composite Fatigue

NASA Technical Reports Server (NTRS)

Chamis, Christos C.

2010-01-01

A multilevel (multiscale/multifunctional) evaluation is demonstrated by applying it to three different sample problems. These problems include the probabilistic evaluation of a space shuttle main engine blade, an engine rotor and an aircraft wing. The results demonstrate that the blade will fail at the highest probability path, the engine two-stage rotor will fail by fracture at the rim and the aircraft wing will fail at 109 fatigue cycles with a probability of 0.9967.

SLS Test Stand Site Selection

NASA Technical Reports Server (NTRS)

Crowe, Kathryn; Williams, Michael

2015-01-01

Test site selection is a critical element of the design, development and production of a new system. With the advent of the new Space Launch System (SLS), the National Aeronautics and Space Administration (NASA) had a number of test site selection decisions that needed to be made early enough in the Program to support the planned Launch Readiness Date (LRD). This case study focuses on decisions that needed to be made in 2011 and 2012 in preparation for the April 2013 DPMC decision about where to execute the Main Propulsion Test that is commonly referred to as "Green Run." Those decisions relied upon cooperative analysis between the Program, the Test Lab and Center Operations. The SLS is a human spaceflight vehicle designed to carry a crew farther into space than humans have previously flown. The vehicle consists of four parts: the crew capsule, the upper stage, the core stage, and the first stage solid rocket boosters. The crew capsule carries the astronauts, while the upper stage, the core stage, and solid rocket boosters provide thrust for the vehicle. In other words, the stages provide the "lift" part of the lift vehicle. In conjunction with the solid rocket boosters, the core stage provides the initial "get-off-the-ground" thrust to the vehicle. The ignition of the four core stage engines and two solid rocket boosters is the first step in the launch portion of the mission. The solid rocket boosters burn out after about 2 minutes of flight, and are then jettisoned. The core stage provides thrust until the vehicle reaches a specific altitude and speed, at which point the core stage is shut off and jettisoned, and the upper stage provides vehicle thrust for subsequent mission trajectories. The integrated core stage primarily consists of a liquid oxygen tank, a liquid hydrogen tank, and the four core stage engines. For the SLS program, four RS-25 engines were selected as the four core stage engines. The RS-25 engine is the same engine that was used for Space Shuttle. The test plan for the integrated core stage was broken down into several segments: Component testing, system level testing, and element level testing. In this context, components are items such as valves, controllers, sensors, etc. Systems are items such as an entire engine, a tank, or the outer stage body. The core stage itself is considered to be an element. The rocket engines are also considered an element. At the program level, it was decided to perform a single green run test on the integrated core stage prior to shipment of it to Kennedy Space Center (KSC) for use in the EM-1 test flight of the SLS vehicle. A green run test is the first live fire of the new integrated core stage and engine elements - without boosters of course. The SLS Program had to decide where to perform SLS green run testing.

The design of two-stage-to-orbit vehicles

NASA Technical Reports Server (NTRS)

1991-01-01

Two separate student design groups developed conceptual designs for a two-stage-to-orbit vehicle, with each design group consisting of a carrier team and an orbiter team. A two-stage-to-orbit system is considered in the event that single-stage-to-orbit is deemed not feasible in the foreseeable future; the two-stage system would also be used as a complement to an already existing heavy lift vehicle. The design specifications given are to lift a 10,000-lb payload 27 ft long by 10 ft diameter, to low Earth orbit (300 n.m.) using an air breathing carrier configuration that will take off horizontally within 15,000 ft. The staging Mach number and altitude were to be determined by the design groups. One group designed a delta wing/body carrier with the orbiter nested within the fuselage of the carrier, and the other group produced a blended cranked-delta wing/body carrier with the orbiter in the more conventional piggyback configuration. Each carrier used liquid hydrogen-fueled turbofanramjet engines, with data provided by General Electric Aircraft Engine Group. While one orbiter used a full-scale Space Shuttle Main Engine (SSME), the other orbiter employed a half-scale SSME coupled with scramjet engines, with data again provided by General Electric. The two groups conceptual designs, along with the technical trade-offs, difficulties, and details that surfaced during the design process are presented.

An Extended Combustion Model for the Aircraft Turbojet Engine

NASA Astrophysics Data System (ADS)

Rotaru, Constantin; Andres-Mihăilă, Mihai; Matei, Pericle Gabriel

2014-08-01

The paper consists in modelling and simulation of the combustion in a turbojet engine in order to find optimal characteristics of the burning process and the optimal shape of combustion chambers. The main focus of this paper is to find a new configuration of the aircraft engine combustion chambers, namely an engine with two main combustion chambers, one on the same position like in classical configuration, between compressor and turbine and the other, placed behind the turbine but not performing the role of the afterburning. This constructive solution could allow a lower engine rotational speed, a lower temperature in front of the first stage of the turbine and the possibility to increase the turbine pressure ratio by extracting the flow stream after turbine in the inner nozzle. Also, a higher thermodynamic cycle efficiency and thrust in comparison to traditional constant-pressure combustion gas turbine engines could be obtained.

CLV First Stage Design, Development, Test and Evaluation

NASA Technical Reports Server (NTRS)

Burt, Richard K.; Brasfield, F.

2006-01-01

The Crew Launch Vehicle (CLV) is an integral part of NASA's Exploration architecture that will provide crew and cargo access to the International Space Station as well as low earth orbit support for lunar missions. Currently in the system definition phase, the CLV is planned to replace the Space Shuttle for crew transport in the post 2010 time frame. It is comprised of a solid rocket booster first stage derived from the current Space Shuttle SRB, a LOX/hydrogen liquid fueled second stage utilizing a derivative of the Space Shuttle Main Engine (SSME) for propulsion, and a Crew Exploration Vehicle (GEV) composed of Command and Service Modules. This paper deals with current DDT&E planning for the CLV first stage solid rocket booster. Described are the current overall point-of-departure design and booster subsystems, systems engineering approach, and milestone schedule requirements.

Multi-Scale/Multi-Functional Probabilistic Composite Fatigue

NASA Technical Reports Server (NTRS)

Chamis, Christos C.

2008-01-01

A multi-level (multi-scale/multi-functional) evaluation is demonstrated by applying it to three different sample problems. These problems include the probabilistic evaluation of a space shuttle main engine blade, an engine rotor and an aircraft wing. The results demonstrate that the blade will fail at the highest probability path, the engine two-stage rotor will fail by fracture at the rim and the aircraft wing will fail at 109 fatigue cycles with a probability of 0.9967.

Design the Cost Approach in Trade-Off's for Structural Components, Illustrated on the Baseline Selection of the Engine Thrust Frame of Ariane 5 ESC-B

NASA Astrophysics Data System (ADS)

Appolloni, L.; Juhls, A.; Rieck, U.

2002-01-01

Designing for value is one of the very actual upcoming methods for design optimization, which broke into the domain of aerospace engineering in the late 90's. In the frame of designing for value two main design philosophies exist: Design For Cost and Design To Cost. Design To Cost is the iterative redesign of a project until the content of the project meets a given budget. Designing For Cost is the conscious use of engineering process technology to reduce life cycle cost while satisfying, and hopefully exceeding, customer demands. The key to understanding cost, and hence to reducing cost, is the ability to measure cost accurately and to allocate it appropriately to products. Only then can intelligent decisions be made. Therefore the necessity of new methods as "Design For Value" or "Design For Competitiveness", set up with a generally multidisciplinary approach to find an optimized technical solution driven by many parameters, depending on the mission scenario and the customer/market needs. Very often three, but not more than five parametric drivers are sufficient. The more variable exist, the higher is in fact the risk to find just a sub-optimized local and not the global optimum, and the less robust is the found solution against change of input parameters. When the main parameters for optimization have been identified, the system engineer has to communicate them to all design engineers, who shall take care of these assessment variables during the entire design and decision process. The design process which has taken to the definition of the feasible structural concepts for the Engine Thrust Frame of the Ariane 5 Upper Cryogenic Stage ESC-B follows these most actual design philosophy methodologies, and combines a design for cost approach, to a design to cost optimization loop. Ariane 5 is the first member of a family of heavy-lift launchers. It aims to evolve into a family of launchers that responds to the space transportation challenges of the 21st century. New upper stages, along with modifications to the main cryogenic stage and solid boosters, will increase performance and meet demands of a changing market. A two-steps approach was decided for future developments of the launcher upper stage, in order to increase the payload lift capability of Ariane 5. The first step ESC-A is scheduled for first launch in 2002. As later step ESC-B shall grow up to 12 tons in GTO orbit, with multiple restart capability, i.e. re-ignitable engine. Ariane 5 ESC-B first flight is targeted for 2006. It will be loaded with 28 metric tons of liquid oxygen and liquid hydrogen and powered by a new expander cycle engine "Vinci". The Vinci engine will be connected to the tanks of the ESC-B stage via the structure named from the designers ETF, or Engine Thrust Frame. In order to develop a design concept for the ETF component a trade off was performed, based on the most modern system engineering methodologies. This paper will describe the basis of the system engineering approach in the design to cost process, and illustrate such approach as it has been applied during the trade off for the baseline selection of the Engine Thrust Frame of Ariane 5 ESC-B.

Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Cockrell, Charles

2008-01-01

NASA is maturing test and evaluation plans leading to flight readiness of the Ares I crew launch vehicle. Key development, qualification, and verification tests are planned . Upper stage engine sea-level and altitude testing. First stage development and qualification motors. Upper stage structural and thermal development and qualification test articles. Main Propulsion Test Article (MPTA). Upper stage green run testing. Integrated Vehicle Ground Vibration Testing (IVGVT). Aerodynamic characterization testing. Test and evaluation supports initial validation flights (Ares I-Y and Orion 1) and design certification.

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 and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

NASA Technical Reports Server (NTRS)

Garcia, C. P.; Medina, C. R.; Protz, C. S.; Kenny, R. J.; Kelly, G. W.; Casiano, M. J.; Hulka, J. R.; Richardson, B. R.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural, thermal, combustion performance, and combustion stability considerations. This paper also discusses the fabrication and assembly of the injector components, including the injector body/interpropellant plate, the additive manufactured GRCop-84 faceplate, and the pieces that make up the injector elements including the oxidizer tube, an inlet to the oxidizer tube, and a facenut that includes the fuel tangential inlets and forms the initial recessed volume where oxidizer and fuel first interact. Hot-fire test results of these main injector designs in an integrated test article that includes an oxidizer-rich preburner are described in companion papers at this JANNAF meeting.

Analyse et design aerodynamique haute-fidelite de l'integration moteur sur un avion BWB

NASA Astrophysics Data System (ADS)

Mirzaei Amirabad, Mojtaba

BWB (Blended Wing Body) is an innovative type of aircraft based on the flying wing concept. In this configuration, the wing and the fuselage are blended together smoothly. BWB offers economical and environmental advantages by reducing fuel consumption through improving aerodynamic performance. In this project, the goal is to improve the aerodynamic performance by optimizing the main body of BWB that comes from conceptual design. The high fidelity methods applied in this project have been less frequently addressed in the literature. This research develops an automatic optimization procedure in order to reduce the drag force on the main body. The optimization is carried out in two main stages: before and after engine installation. Our objective is to minimize the drag by taking into account several constraints in high fidelity optimization. The commercial software, Isight is chosen as an optimizer in which MATLAB software is called to start the optimization process. Geometry is generated using ANSYS-DesignModeler, unstructured mesh is created by ANSYS-Mesh and CFD calculations are done with the help of ANSYS-Fluent. All of these software are coupled together in ANSYS-Workbench environment which is called by MATLAB. The high fidelity methods are used during optimization by solving Navier-Stokes equations. For verifying the results, a finer structured mesh is created by ICEM software to be used in each stage of optimization. The first stage includes a 3D optimization on the surface of the main body, before adding the engine. The optimized case is then used as an input for the second stage in which the nacelle is added. It could be concluded that this study leads us to obtain appropriate reduction in drag coefficient for BWB without nacelle. In the second stage (adding the nacelle) a drag minimization is also achieved by performing a local optimization. Furthermore, the flow separation, created in the main body-nacelle zone, is reduced.

Low-thrust chemical propulsion system pump technology

NASA Technical Reports Server (NTRS)

Sabiers, R. L.; Siebenhaar, A.

1981-01-01

Candidate pump and driver systems for low thrust cargo orbit transfer vehicle engines which deliver large space structures to geosynchronous equatorial orbit and beyond are evaluated. The pumps operate to 68 atmospheres (1000 psi) discharge pressure and flowrates suited to cryogenic engines using either LOX/methane or LOX/hydrogen propellants in thrust ranges from 445 to 8900 N (100 to 2000 lb F). Analysis of the various pumps and drivers indicate that the low specific speed requirement will make high fluid efficiencies difficult to achieve. As such, multiple stages are required. In addition, all pumps require inducer stages. The most attractive main pumps are the multistage centrifugal pumps.

Reduced Order Modeling of SLS Liquid Hydrogen Pre-Valve Flow Guide to Enable Rapid Transient Analysis

NASA Technical Reports Server (NTRS)

Brown, Andrew M.; Mulder, Andrew

2017-01-01

NASA is developing a new launch vehicle, called the Space Launch System (SLS), which is intended on taking humans out of low earth orbit to destinations including the moon, asteroids, and Mars. The propulsion system for the core stage of this vehicle includes four RS-25 Liquid Hydrogen/Oxygen rocket engines. These engines are upgraded versions of the Space Shuttle Main Engines (SSME); the upgrades include higher power levels and affordability enhancements. As with any new vehicle, the Main Propulsion System (MPS), which include the feedlines and ancillary hardware connecting the engines to the fuel and oxidizer tanks, had to be redesigned (figure 1 - export clearance in progress), as the previous MPS for the SSME's was inherently part of the Space Shuttle System, which had a completely different overall configuration.

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.

Assessment of crack growth in a space shuttle main engine first-stage, high-pressure fuel turbopump blade

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali

1993-01-01

A two-dimensional finite element fracture mechanics analysis of a space shuttle main engine (SSME) turbine blade firtree was performed using the MARC finite element code. The analysis was conducted under combined effects of thermal and mechanical loads at steady-state conditions. Data from a typical engine stand cycle of the SSME were used to run a heat transfer analysis and, subsequently, a thermal structural fracture mechanics analysis. Temperature and stress contours for the firtree under these operating conditions were generated. High stresses were found at the firtree lobes where crack initiation was triggered. A life assessment of the firtree was done by assuming an initial and a final crack size.

Orbit transfer rocket engine technology program

NASA Technical Reports Server (NTRS)

Gustafson, N. B.; Harmon, T. J.

1993-01-01

An advanced near term (1990's) space-based Orbit Transfer Vehicle Engine (OTVE) system was designed, and the technologies applicable to its construction, maintenance, and operations were developed under Tasks A through F of the Orbit Transfer Rocket Engine Technology Program. Task A was a reporting task. In Task B, promising OTV turbomachinery technologies were explored: two stage partial admission turbines, high velocity ratio diffusing crossovers, soft wear ring seals, advanced bearing concepts, and a rotordynamic analysis. In Task C, a ribbed combustor design was developed. Possible rib and channel geometries were chosen analytically. Rib candidates were hot air tested and laser velocimeter boundary layer analyses were conducted. A channel geometry was also chosen on the basis of laser velocimeter data. To verify the predicted heat enhancement effects, a ribbed calorimeter spool was hot fire tested. Under Task D, the optimum expander cycle engine thrust, performance and envelope were established for a set of OTV missions. Optimal nozzle contours and quick disconnects for modularity were developed. Failure Modes and Effects Analyses, maintenance and reliability studies and component study results were incorporated into the engine system. Parametric trades on engine thrust, mixture ratio, and area ratio were also generated. A control system and the health monitoring and maintenance operations necessary for a space-based engine were outlined in Task E. In addition, combustor wall thickness measuring devices and a fiberoptic shaft monitor were developed. These monitoring devices were incorporated into preflight engine readiness checkout procedures. In Task F, the Integrated Component Evaluator (I.C.E.) was used to demonstrate performance and operational characteristics of an advanced expander cycle engine system and its component technologies. Sub-system checkouts and a system blowdown were performed. Short transitions were then made into main combustor ignition and main stage operation.

NASA's J-2X Engine Builds on the Apollo Program for Lunar Return Missions

NASA Technical Reports Server (NTRS)

Snoddy, Jimmy R.

2006-01-01

In January 2006, NASA streamlined its U.S. Vision for Space Exploration hardware development approach for replacing the Space Shuttle after it is retired in 2010. The revised CLV upper stage will use the J-2X engine, a derivative of NASA s Apollo Program Saturn V s S-II and S-IVB main propulsion, which will also serve as the Earth Departure Stage (EDS) engine. This paper gives details of how the J- 2X engine effort mitigates risk by building on the Apollo Program and other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with first demonstration flight in 2010 and human test flights in 2012. It is well documented that propulsion is historically a high-risk area. NASA s risk reduction strategy for the J-2X engine design, development, test, and evaluation is to build upon heritage hardware and apply valuable experience gained from past development efforts. In addition, NASA and its industry partner, Rocketdyne, which originally built the J-2, have tapped into their extensive databases and are applying lessons conveyed firsthand by Apollo-era veterans of America s first round of Moon missions in the 1960s and 1970s. NASA s development approach for the J-2X engine includes early requirements definition and management; designing-in lessons learned from the 5-2 heritage programs; initiating long-lead procurement items before Preliminary Desi& Review; incorporating design features for anticipated EDS requirements; identifying facilities for sea-level and altitude testing; and starting ground support equipment and logistics planning at an early stage. Other risk reduction strategies include utilizing a proven gas generator cycle with recent development experience; utilizing existing turbomachinery ; applying current and recent main combustion chamber (Integrated Powerhead Demonstrator) and channel wall nozzle (COBRA) advances; and performing rigorous development, qualification, and certification testing of the engine system, with a philosophy of "test what you fly, and fly what you test". These and other active risk management strategies are in place to deliver the J-2X engine for LEO and lunar return missions as outlined in the U.S. Vision for Space Exploration.

Smoke and fire Rocket-engine ablaze on This Week @NASA – August 14, 2015

NASA Image and Video Library

2015-08-14

On Aug. 13, NASA conducted a test firing of the RS-25 rocket engine at Stennis Space Center. The 535 second test was the sixth in the current series of seven developmental tests of the former space shuttle main engine. Four RS-25 engines will power the core stage of the new Space Launch System (SLS) rocket, which will carry humans deeper into space than ever before, including to an asteroid and Mars. Also, Veggies in space, Russian spacewalk, Supply ship undocks from ISS, Smallest giant black hole, 10th anniversary of MRO launch and more!

Propulsion System Advances that Enable a Reusable Liquid Fly Back Booster (LFBB)

NASA Technical Reports Server (NTRS)

Keith, Edward L.; Rothschild, William J.

1998-01-01

This paper provides an overview of the booster propulsion system for the Liquid Fly Back Booster (LFBB). This includes, system requirements, design approach, concept of operations, reliability, safety and cost assumptions. The paper summarizes the findings of the Boeing propulsion team that has been studying the LFBB feasibility as a booster replacement for the Space Shuttle. This paper will discuss recent advances including a new generation of kerosene and oxygen rich pre-burner staged combustion cycle main rocket engines. The engine reliability and safety is expected to be much higher than current standards by adding extra operating margins into the design and normally operating the engines at 75% of engine rated power. This allows for engine out capability. The new generation of main engines operates at significantly higher chamber pressure than the prior generation of gas generator cycle engines. The oxygen rich pre-burner engine cycle, unlike the fuel rich gas generator cycle, results in internally self-cleaning firings which facilitates reusability. Maintenance is further enhanced with integrated health monitoring to improve safety and turn-around efficiency. The maintainability of the LFBB LOX / kerosene engines is being improved by designing the vehicle/engine interfaces for easy access to key engine components.

Propulsion system advances that enable a reusable Liquid Fly Back Booster (LFBB)

NASA Technical Reports Server (NTRS)

Keith, E. L.; Rothschild, W. J.

1998-01-01

This paper provides an overview of the booster propulsion system for the Liquid Fly Back Booster (LFBB). This includes, system requirements, design approach, concept of operations, reliability, safety and cost assumptions. The paper summarizes the findings of the Boeing propulsion team that has been studying the LFBB feasibility as a booster replacement for the Space Shuttle. This paper will discuss recent advances including a new generation of kerosene and oxygen rich pre-burner staged combustion cycle main rocket engines. The engine reliability and safety is expected to be much higher than current standards by adding extra operating margins into the design and normally operating the engines at 75% of engine rated power. This allows for engine out capability. The new generation of main engines operates at significantly higher chamber pressure than the prior generation of gas generator cycle engines. The oxygen rich pre-burner engine cycle, unlike the fuel rich gas generator cycle, results in internally self-cleaning firings which facilitates reusability. Maintenance is further enhanced with integrated health monitoring to improve safety and turn-around efficiency. The maintainability of the LFBB LOX/kerosene engines is being improved by designing the vehicle/engine interfaces for easy access to key engine components.

Rocket nozzle expansion ratio analysis for dual-fuel earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Martin, James A.

1989-01-01

Results are reported from a recent study of the effects of Space Shuttle Main Engine expansion ratio modifications, in the cases of both single-stage and two-stage systems. Two-position nozzles were employed; after varying the lower expansion ratio while the higher was held constant at 120, the lower expansion ratio was held constant at 40 or 60 while the higher expansion ratio was varied. The expansion ratios for minimum vehicle dry mass are different for single-stage and two-stage systems. For two-stage systems, a single expansion ratio of 77.5 provides a lower dry mass than any two-position nozzle.

Engine Evaluation of Advanced Technology Control Components

DTIC Science & Technology

1976-08-01

producer turbine rotor blades. This is a very desirable control feature, because protecting turbine blades from overtemperature is particularly...centrifugal boost stage operating back to back on a common drive shaft that is direct driven through the alter- nator rotor shaft. The main stage is an...computation makes this simple dynamic pumping machine possible. The pen- alty of this simple design is lower overall efficiency as com- pared to a

System Noise Assessment of Blended-Wing-Body Aircraft With Open Rotor Propulsion

NASA Technical Reports Server (NTRS)

Guo, Yueping; Thomas, Russell H.

2015-01-01

An aircraft system noise study is presented for the Blended-Wing-Body (BWB) aircraft concept with three open rotor engines mounted on the upper surface of the airframe. It is shown that for such an aircraft, the cumulative Effective Perceived Noise Level (EPNL) is about 24 dB below the current aircraft noise regulations of Stage 4. While this makes the design acoustically viable in meeting the regulatory requirements, even with the consideration of more stringent noise regulations of a possible Stage 5 in the next decade or so, the design will likely meet stiff competitions from aircraft with turbofan engines. It is shown that the noise levels of the BWB design are held up by the inherently high noise levels of the open rotor engines and the limitation on the shielding benefit due to the practical design constraint on the engine location. Furthermore, it is shown that the BWB design has high levels of noise from the main landing gear, due to their exposure to high speed flow at the junction between the center body and outer wing. These are also the reasons why this baseline BWB design does not meet the NASA N+2 noise goal of 42 dB below Stage 4. To identify approaches that may further reduce noise, parametric studies are also presented, including variations in engine location, vertical tail and elevon variations, and airframe surface acoustic liner treatment effect. These have the potential to further reduce noise but they are only at the conceptual stage.

Detection of frequency-mode-shift during thermoacoustic combustion oscillations in a staged aircraft engine model combustor

NASA Astrophysics Data System (ADS)

Kobayashi, Hiroaki; Gotoda, Hiroshi; Tachibana, Shigeru; Yoshida, Seiji

2017-12-01

We conduct an experimental study using time series analysis based on symbolic dynamics to detect a precursor of frequency-mode-shift during thermoacoustic combustion oscillations in a staged aircraft engine model combustor. With increasing amount of the main fuel, a significant shift in the dominant frequency-mode occurs in noisy periodic dynamics, leading to a notable increase in oscillation amplitudes. The sustainment of noisy periodic dynamics during thermoacoustic combustion oscillations is clearly shown by the multiscale complexity-entropy causality plane in terms of statistical complexity. A modified version of the permutation entropy allows us to detect a precursor of the frequency-mode-shift before the amplification of pressure fluctuations.

National Launch System Space Transportation Main Engine

NASA Technical Reports Server (NTRS)

Hoodless, Ralph M., Jr.; Monk, Jan C.; Cikanek, Harry A., III

1991-01-01

The present liquid-oxygen/liquid-hydrogen engine is described as meeting the specific requirements of the National Launch System (NLS) Program including cost-effectiveness and robustness. An overview of the NLS and its objectives is given which indicates that the program aims to develop a flexible launch system to meet security, civil, and commercial needs. The Space Transportation Main Engine (STME) provides core and boost propulsion for the 1.5-stage vehicle and core propulsion for the solid booster vehicle. The design incorporates step-throttling, order-of-magnitude reductions in welds, and configuration targets designed to optimize robustness. The STME is designed to provide adaptable and dependable propulsion while minimizing recurring costs and is designed to meet the needs of NLS and other typical space-transportation programs currently being planned.

Power processor for a 20CM ion thruster

NASA Technical Reports Server (NTRS)

Biess, J. J.; Schoenfeld, A. D.; Cohen, E.

1973-01-01

A power processor breadboard for the JPL 20CM Ion Engine was designed, fabricated, and tested to determine compliance with the electrical specification. The power processor breadboard used the silicon-controlled rectifier (SCR) series resonant inverter as the basic power stage to process all the power to the ion engine. The breadboard power processor was integrated with the JPL 20CM ion engine and complete testing was performed. The integration tests were performed without any silicon-controlled rectifier failure. This demonstrated the ruggedness of the series resonant inverter in protecting the switching elements during arcing in the ion engine. A method of fault clearing the ion engine and returning back to normal operation without elaborate sequencing and timing control logic was evolved. In this method, the main vaporizer was turned off and the discharge current limit was reduced when an overload existed on the screen/accelerator supply. After the high voltage returned to normal, both the main vaporizer and the discharge were returned to normal.

Application of fault factor method to fault detection and diagnosis for space shuttle main engine

NASA Astrophysics Data System (ADS)

Cha, Jihyoung; Ha, Chulsu; Ko, Sangho; Koo, Jaye

2016-09-01

This paper deals with an application of the multiple linear regression algorithm to fault detection and diagnosis for the space shuttle main engine (SSME) during a steady state. In order to develop the algorithm, the energy balance equations, which balances the relation among pressure, mass flow rate and power at various locations within the SSME, are obtained. Then using the measurement data of some important parameters of the engine, fault factors which reflects the deviation of each equation from the normal state are estimated. The probable location of each fault and the levels of severity can be obtained from the estimated fault factors. This process is numerically demonstrated for the SSME at 104% Rated Propulsion Level (RPL) by using the simulated measurement data from the mathematical models of the engine. The result of the current study is particularly important considering that the recently developed reusable Liquid Rocket Engines (LREs) have staged-combustion cycles similarly to the SSME.

Analysis of the Functionality of Refillable Propellant Management Devices (PMD)

NASA Astrophysics Data System (ADS)

Winkelmann, Yvonne; Gaulke, Diana; Dreyer, Michael E.

In order to restart a stage of a spacecraft it is necessary to position the liquid stable over the tank outlet. The gas-or vapor-free provision of the thrusters for the main engine start-up can be accomplished by the use of propellant management devices (PMDs). A propellant refillable reservoir (PRR) will supply the engine with the required amount of liquid propellant until the liquid outside the PRR has settled at the bottom of the tank. Hence, the reservoir will be refilled and the main engine can be restarted. This technique has been applied in case of storable propellants yet, e.g. in satellites or ATVs. For the application in a cryogenic upper stage demonstration and validation tests are still necessary. Ground experiments to simulate propulsed phases are evaluated. To demonstrate the functionality under propulsed conditions first filling, draining and draining with a constant fill level of the tank (refilling) are analyzed. Different inflows with respect to filling and varied outflow rates for the draining tests are investigated. Pressure losses in the LOX-PMD are measured during draining and compared to a previously accomplished estimation with an one-dimensional streamtube theory.

Space Shuttle Main Engine High Pressure Fuel Turbopump Turbine Blade Cracking

NASA Technical Reports Server (NTRS)

Lee, Henry

1988-01-01

The analytical results from two-dimensional (2D) and three-dimensional (3D) finite element model investigations into the cracking of Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) first- and second-stage turbine blades are presented. Specifically, the initiation causes for transverse cracks on the pressure side of the firststage blade fir tree lobes and face/corner cracks on the downstream fir tree face of the second-state blade are evaluated. Because the blade material, MAR-M-246 Hf (DS), is highly susceptible to hydrogen embrittlement in the -100 F to 400 F thermal environment, a steady-state condition (full power level = 109 percent) rather than a start-up or shut-down transient was considered to be the most likely candidate for generating a high-strain state in the fir tree areas. Results of the analyses yielded strain levels on both first- and second-stage blade fir tree regions that are of a magnitude to cause hydrogen assisted low cycle fatigue cracking. Also evident from the analysis is that a positive margin against fir tree cracking exists for the planned design modifications, which include shot peening for both first- and second-stage blade fir tree areas.

The J-2X Upper Stage Engine: From Heritage to Hardware

NASA Technical Reports Server (NTRS)

Byrd, THomas

2008-01-01

NASA's Global Exploration Strategy requires safe, reliable, robust, efficient transportation to support sustainable operations from Earth to orbit and into the far reaches of the solar system. NASA selected the Ares I crew launch vehicle and the Ares V cargo launch vehicle to provide that transportation. Guiding principles in creating the architecture represented by the Ares vehicles were the maximum use of heritage hardware and legacy knowledge, particularly Space Shuttle assets, and commonality between the Ares vehicles where possible to streamline the hardware development approach and reduce programmatic, technical, and budget risks. The J-2X exemplifies those goals. It was selected by the Exploration Systems Architecture Study (ESAS) as the upper stage propulsion for the Ares I Upper Stage and the Ares V Earth Departure Stage (EDS). The J-2X is an evolved version ofthe historic J-2 engine that successfully powered the second stage of the Saturn I launch vehicle and the second and third stages of the Saturn V launch vehicle. The Constellation architecture, however, requires performance greater than its predecessor. The new architecture calls for larger payloads delivered to the Moon and demands greater loss of mission reliability and numerous other requirements associated with human rating that were not applied to the original J-2. As a result, the J-2X must operate at much higher temperatures, pressures, and flow rates than the heritage J-2, making it one of the highest performing gas generator cycle engines ever built, approaching the efficiency of more complex stage combustion engines. Development is focused on early risk mitigation, component and subassembly test, and engine system test. The development plans include testing engine components, including the subscale injector, main igniter, powerpack assembly (turbopumps, gas generator and associated ducting and structural mounts), full-scale gas generator, valves, and control software with hardware-in-the-loop. Testing expanded in 2007, accompanied by the refinement of the design through several key milestones. This paper discusses those 2007 tests and milestones, as well as updates key developments in 2008.

Development of Implantable Medical Devices: From an Engineering Perspective

PubMed Central

2013-01-01

From the first pacemaker implant in 1958, numerous engineering and medical activities for implantable medical device development have faced challenges in materials, battery power, functionality, electrical power consumption, size shrinkage, system delivery, and wireless communication. With explosive advances in scientific and engineering technology, many implantable medical devices such as the pacemaker, cochlear implant, and real-time blood pressure sensors have been developed and improved. This trend of progress in medical devices will continue because of the coming super-aged society, which will result in more consumers for the devices. The inner body is a special space filled with electrical, chemical, mechanical, and marine-salted reactions. Therefore, electrical connectivity and communication, corrosion, robustness, and hermeticity are key factors to be considered during the development stage. The main participants in the development stage are the user, the medical staff, and the engineer or technician. Thus, there are three different viewpoints in the development of implantable devices. In this review paper, considerations in the development of implantable medical devices will be presented from the viewpoint of an engineering mind. PMID:24143287

Integrated System Test Approaches for the NASA Ares I Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.; Askins, Bruce R.; Bland, Jeffrey; Davis, Stephan; Holladay, Jon B.; Taylor, James L.; Taylor, Terry L.; Robinson, Kimberly F.; Roberts, Ryan E.; Tuma, Margaret

2007-01-01

The Ares I Crew Launch Vehicle (CLV) is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew access to the International Space Station (ISS) and, together with the Ares V Cargo Launch Vehicle (CaLV), serves as one component of a future launch capability for human exploration of the Moon. During the system requirements definition process and early design cycles, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements: the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine, will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle dynamic test (IVDT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, and a highaltitude actuation of the launch abort system (LAS) following separation. The Orion-1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.

Ascent performance feasibility for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

Probabilistic structural analysis methods for space transportation propulsion systems

NASA Technical Reports Server (NTRS)

Chamis, C. C.; Moore, N.; Anis, C.; Newell, J.; Nagpal, V.; Singhal, S.

1991-01-01

Information on probabilistic structural analysis methods for space propulsion systems is given in viewgraph form. Information is given on deterministic certification methods, probability of failure, component response analysis, stress responses for 2nd stage turbine blades, Space Shuttle Main Engine (SSME) structural durability, and program plans. .

Application of Weibull analysis to SSME hardware

NASA Technical Reports Server (NTRS)

Gray, L. A. B.

1986-01-01

Generally, it has been documented that the wearing of engine parts forms a failure distribution which can be approximated by a function developed by Weibull. The purpose here is to examine to what extent the Weibull distribution approximates failure data for designated engine parts of the Space Shuttle Main Engine (SSME). The current testing certification requirements will be examined in order to establish confidence levels. An examination of the failure history of SSME parts/assemblies (turbine blades, main combustion chamber, or high pressure fuel pump first stage impellers) which are limited in usage by time or starts will be done by using updated Weibull techniques. Efforts will be made by the investigator to predict failure trends by using Weibull techniques for SSME parts (turbine temperature sensors, chamber pressure transducers, actuators, and controllers) which are not severely limited by time or starts.

Extended duration lunar lander

NASA Technical Reports Server (NTRS)

Babic, Nikola; Carter, Matt; Cosper, Donna; Garza, David; Gonzalez, Eloy; Goodine, David; Hirst, Edward; Li, Ray; Lindsey, Martin; Ng, Tony

1993-01-01

Selenium Technologies has been conducting preliminary design work on a manned lunar lander for use in NASA's First Lunar Outpost (FLO) program. The resulting lander is designed to carry a crew of four astronauts to a prepositioned habitat on the lunar surface, remain on the lunar surface for up to 45 days while the crew is living in the habitat, then return the crew to earth via direct reentry and land recovery. Should the need arise, the crew can manually guide the lander to a safe lunar landing site, and live in the lander for up to ten days on the surface. Also, an abort to earth is available during any segment of the mission. The main propulsion system consists of a cluster of four modified Pratt and Whitney RL10 rocket engines that use liquid methane (LCH4) and liquid oxygen (LOX). Four engines are used to provide redundancy and a satisfactory engine out capability. Differences between the new propulsion system and the original system include slightly smaller engine size and lower thrust per engine, although specific impulse remains the same despite the smaller size. Concerns over nozzle ground clearance and engine reliability, as well as more information from Pratt and Whitney, brought about this change. The power system consists of a combination of regenerative fuel cells and solar arrays. While the lander is in flight to or from the moon, or during the lunar night, fuel cells provide all electrical power. During the lunar day, solar arrays are deployed to provide electrical power for the lander as well as electrolyzers, which separate some water back into hydrogen and oxygen for later use by the fuel cells. Total storage requirements for oxygen, hydrogen, and water are 61 kg, 551 kg, and 360 kg, respectively. The lander is a stage-and-a-half design with descent propellant, cargo, and landing gear contained in the descent stage, and the main propulsion system, ascent propellant, and crew module contained in the ascent stage. The primary structure for both stages is a truss, to which all tanks and components are attached. The crew module is a conical shape similar to that of the Apollo Command Module, but significantly larger with a height and maximum diameter of six meters.

An analysis of the booster plume impingement environment during the space shuttle nominal staging maneuver

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.; Penny, M. M.; Greenwood, T. F.; Fossler, I. H.

1972-01-01

An experimental study of the plume impingement heating on the space shuttle booster afterbody resulting from the space shuttle orbiter engine plumes was conducted. The 1/100-scale model tests consisted of one and two orbiter engine firings on a flat plate, a flat plate with a fin, and a cylinder model. The plume impingement heating rates on these surfaces were measured using thin film heat transfer gages. Results indicate the engine simulation is a reasonable approximation to the two engine configuration, but more tests are needed to verify the plume model of the main engine configuration. For impingment, results show models experienced laminar boundary layer convective heating. Therefore, tests at higher Reynolds numbers are needed to determine impingment heating.

NASA Marches on with Test of RS-25 Engine for New Space Launch System

NASA Image and Video Library

2016-07-29

NASA engineers conducted a successful developmental test of RS-25 rocket engine No. 0528 July 29, 2016, to collect critical performance data for the most powerful rocket in the world – the Space Launch System (SLS). The engine roared to life for a full 650-second test on the A-1 Test Stand at NASA’s Stennis Space Center, near Bay St. Louis, Mississippi, marking another step forward in development of the SLS, which will launch humans deeper into space than ever before, including on the journey to Mars. Four RS-25 engines, joined with a pair of solid rocket boosters, will power the SLS core stage at launch. The RS-25 engines used on the first four SLS flights are former space shuttle main engines, modified to operate at a higher performance level and with a new engine controller, which allows communication between the vehicle and engine.

Emission Reduction of Fuel-Staged Aircraft Engine Combustor Using an Additional Premixed Fuel Nozzle.

PubMed

Yamamoto, Takeshi; Shimodaira, Kazuo; Yoshida, Seiji; Kurosawa, Yoji

2013-03-01

The Japan Aerospace Exploration Agency (JAXA) is conducting research and development on aircraft engine technologies to reduce environmental impact for the Technology Development Project for Clean Engines (TechCLEAN). As a part of the project, combustion technologies have been developed with an aggressive target that is an 80% reduction over the NO x threshold of the International Civil Aviation Organization (ICAO) Committee on Aviation Environmental Protection (CAEP)/4 standard. A staged fuel nozzle with a pilot mixer and a main mixer was developed and tested using a single-sector combustor under the target engine's landing and takeoff (LTO) cycle conditions with a rated output of 40 kN and an overall pressure ratio of 25.8. The test results showed a 77% reduction over the CAEP/4 NO x standard. However, the reduction in smoke at thrust conditions higher than the 30% MTO condition and of CO emission at thrust conditions lower than the 85% MTO condition are necessary. In the present study, an additional fuel burner was designed and tested with the staged fuel nozzle in a single-sector combustor to control emissions. The test results show that the combustor enables an 82% reduction in NO x emissions relative to the ICAO CAEP/4 standard and a drastic reduction in smoke and CO emissions.

Staged combustion with piston engine and turbine engine supercharger

DOEpatents

Fischer, Larry E [Los Gatos, CA; Anderson, Brian L [Lodi, CA; O'Brien, Kevin C [San Ramon, CA

2006-05-09

A combustion engine method and system provides increased fuel efficiency and reduces polluting exhaust emissions by burning fuel in a two-stage combustion system. Fuel is combusted in a piston engine in a first stage producing piston engine exhaust gases. Fuel contained in the piston engine exhaust gases is combusted in a second stage turbine engine. Turbine engine exhaust gases are used to supercharge the piston engine.

Staged combustion with piston engine and turbine engine supercharger

DOEpatents

Fischer, Larry E [Los Gatos, CA; Anderson, Brian L [Lodi, CA; O'Brien, Kevin C [San Ramon, CA

2011-11-01

A combustion engine method and system provides increased fuel efficiency and reduces polluting exhaust emissions by burning fuel in a two-stage combustion system. Fuel is combusted in a piston engine in a first stage producing piston engine exhaust gases. Fuel contained in the piston engine exhaust gases is combusted in a second stage turbine engine. Turbine engine exhaust gases are used to supercharge the piston engine.

Design for Production Manual. Volume 3. The Application of Production Engineering

DTIC Science & Technology

1985-12-01

3.1.1.1 Product Development The main objective of a product development strategy is to develop desiqns which are both highly marketable and which-allow the...responsibility between stages. Control is exercised by workers and foremen at each stage. 3-1 /104 3.1.1.2 Marketing and sales As a result of a well defined...to become even more competitive by achieving greater overlap, even for a mix of ships being constructed simultaneously. Their goal is to achieve

Space shuttle main engine: Interactive design challenges

NASA Technical Reports Server (NTRS)

Mccarty, J. P.; Wood, B. K.

1985-01-01

The operating requirements established by NASA for the SSME were considerably more demanding than those for earlier rocket engines used in the military launch vehicles or Apollo program. The SSME, in order to achieve the high performance, low weight, long life, reusable objectives, embodied technical demands far in excess of its predecessor rocket engines. The requirements dictated the use of high combustion pressure and the staged combustion cycle which maximizes performance through total use of all propellants in the main combustion process. This approach presented a myriad of technical challenges for maximization of performance within attainable state of the art capabilities for operating pressures, operating temperatures and rotating machinery efficiencies. Controlling uniformity of the high pressure turbomachinery turbine temperature environment was a key challenge for thrust level and life capability demanding innovative engineering. New approaches in the design of the components were necessary to accommodate the multiple use, minimum maintenance objectives. Included were the use of line replaceable units to facilitate field maintenance automatic checkout and internal inspection capabilities.

Tissue engineering as innovative chance for organ replacement in radical tumor surgery.

PubMed

Alberti, C

2013-03-01

Different pathological conditions such as congenital organ absence, severe organ injuries, end-stage organ failure and malignancy-related organ removal, have few effective therapeutic options a part from a whole organ transplant, that, however, often meets with a serious shortage of suitable donor organs. The purpose of this paper consists in highlighting what the novel tissue engineering approaches might help to solve such problems. EMERGING CONCEPTS: A recent approach in tissue/organ engineering, particularly to build bioartificial airways, is the procedure of decellularizing a whole donor organ to obtain a complex 3D-biomatrix-scaffold maintaining the intrinsic vascular network, that is subsequently recellularized with recipient's autologous organ-specific differentiated cells or/and stem cells, to build a potentially functional biological substitute. Such strategy has been clinically used to replace organ in trachea/broncus tumor patients. In another approach, mainly used to construct a bioartificial urinary bladder tissue, different types of either biodegradable synthetic polymers or naturally-derived matrices or even polymer/biomatrix-composite materials are used as scaffold for either cell-free or autologous cell-seeded tissue engineering procedures. So far, such technique has been mainly used to make an augmentation cystoplasty in patients with end-stage poorly compliant neuropathic bladder or in exstrophic bladder subjects. Intriguing developments in biomaterial science, nanotechnologies, stem cell biology, and further improvements in bioreactor manufacturing will allow to generate, in the near future, tissue engineered organs that, as for structure/function so the native one-like, might represent the optimum solution to replace organs in tumor surgery.

The Control System for the X-33 Linear Aerospike Engine

NASA Technical Reports Server (NTRS)

Jackson, Jerry E.; Espenschied, Erich; Klop, Jeffrey

1998-01-01

The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

Resistance of Metallic Screens in a Cryogenic Flow

NASA Astrophysics Data System (ADS)

Fischer, Alexander; Stief, Malte

The propellant behaviour in cryogenic upper stages tanks imposes challenging requirements on the design, especially for future upper stages designed for multiple restarts and long ballistic flight phases. The main challenge is the supply of the propellants to the feed system prior to the engine reignition. During the entire mission the engine requires a gaseous and bubble free liquid supply of propellant at the required thermodynamic conditions. The current research focus is to prepare the initial steps for the maturation of the Propellant Management Device (PMD) technology for cryogenic tank systems. Main components of such a PMD are metallic screens. The metallic screens are used as barrier for any gas bubbles within the fluid stream approaching the space craft engines. The screen characteristics are of fundamental importance for the PMD and feed system design. The paper presents a summary on available experimental screen data with regard to the flow resistance and gives a comparison with theoretical and empirical predictions found in literature. The lack on comparable data with regard to space craft applications and the need on further research with cryogenic flows is demonstrated. The DLR Institute of Space Systems is preparing various cryogenic tests to collect the desired information about the flow properties of such metallic screens. The planned test setup and the foreseen experiments will be presented.

Marshall Space Flight Center CFD overview

NASA Technical Reports Server (NTRS)

Schutzenhofer, Luke A.

1989-01-01

Computational Fluid Dynamics (CFD) activities at Marshall Space Flight Center (MSFC) have been focused on hardware specific and research applications with strong emphasis upon benchmark validation. The purpose here is to provide insight into the MSFC CFD related goals, objectives, current hardware related CFD activities, propulsion CFD research efforts and validation program, future near-term CFD hardware related programs, and CFD expectations. The current hardware programs where CFD has been successfully applied are the Space Shuttle Main Engines (SSME), Alternate Turbopump Development (ATD), and Aeroassist Flight Experiment (AFE). For the future near-term CFD hardware related activities, plans are being developed that address the implementation of CFD into the early design stages of the Space Transportation Main Engine (STME), Space Transportation Booster Engine (STBE), and the Environmental Control and Life Support System (ECLSS) for the Space Station. Finally, CFD expectations in the design environment will be delineated.

Reefing Line Tension in CPAS Main Parachute Clusters

NASA Technical Reports Server (NTRS)

Ray, Eric S.

2013-01-01

Reefing lines are an essential feature to manage inflation loads. During each Engineering Development Unit (EDU) test of the Capsule Parachute Assembly System (CPAS), a chase aircraft is staged to be level with the cluster of Main ringsail parachutes during the initial inflation and reefed stages. This allows for capturing high-quality still photographs of the reefed skirt, suspension line, and canopy geometry. The over-inflation angles are synchronized with measured loads data in order to compute the tension force in the reefing line. The traditional reefing tension equation assumes radial symmetry, but cluster effects cause the reefed skirt of each parachute to elongate to a more elliptical shape. This effect was considered in evaluating multiple parachutes to estimate the semi-major and semi-minor axes. Three flight tests are assessed, including one with a skipped first stage, which had peak reefing line tension over three times higher than the nominal parachute disreef sequence.

Investigation of the effect of pilot burner on lean blow out performance of a staged injector

NASA Astrophysics Data System (ADS)

Yang, Jinhu; Zhang, Kaiyu; Liu, Cunxi; Ruan, Changlong; Liu, Fuqiang; Xu, Gang

2014-12-01

The staged injector has exhibited great potential to achieve low emissions and is becoming the preferable choice of many civil airplanes. Moreover, it is promising to employ this injector design in military engine, which requires most of the combustion air enters the combustor through injector to reduce smoke emission. However, lean staged injector is prone to combustion instability and extinction in low load operation, so techniques for broadening its stable operation ranges are crucial for its application in real engine. In this work, the LBO performance of a staged injector is assessed and analyzed on a single sector test section. The experiment was done in atmospheric environment with optical access. Kerosene-PLIF technique was used to visualize the spray distribution and common camera was used to record the flame patterns. Emphasis is put on the influence of pilot burner on LBO performance. The fuel to air ratios at LBO of six injectors with different pilot swirler vane angle were evaluated and the obtained LBO data was converted into data at idle condition. Results show that the increase of pilot swirler vane angle could promote the air assisted atomization, which in turn improves the LBO performance slightly. Flame patterns typical in the process of LBO are analyzed and attempts are made to find out the main factors which govern the extinction process with the assistance of spray distribution and numerical flow field results. It can be learned that the flame patterns are mainly influenced by structure of the flow field just behind the pilot burner when the fuel mass flow rate is high; with the reduction of fuel, atomization quality become more and more important and is the main contributing factor of LBO. In the end of the paper, conclusions are drawn and suggestions are made for the optimization of the present staged injector.

Statistical analysis of 59 inspected SSME HPFTP turbine blades (uncracked and cracked)

NASA Technical Reports Server (NTRS)

Wheeler, John T.

1987-01-01

The numerical results of statistical analysis of the test data of Space Shuttle Main Engine high pressure fuel turbopump second-stage turbine blades, including some with cracks are presented. Several statistical methods use the test data to determine the application of differences in frequency variations between the uncracked and cracked blades.

A transient model of the RL10A-3-3A rocket engine

NASA Technical Reports Server (NTRS)

Binder, Michael P.

1995-01-01

RL10A-3-3A rocket engines have served as the main propulsion system for Centaur upper stage vehicles since the early 1980's. This hydrogen/oxygen expander cycle engine continues to play a major role in the American launch industry. The Space Propulsion Technology Division at the NASA Lewis Research Center has created a computer model of the RL10 engine, based on detailed component analyses and available test data. This RL10 engine model can predict the performance of the engine over a wide range of operating conditions. The model may also be used to predict the effects of any proposed design changes and anticipated failure scenarios. In this paper, the results of the component analyses are discussed. Simulation results from the new system model are compared with engine test and flight data, including the start and shut-down transient characteristics.

Tip-Clearance Measurement in the First Stage of the Compressor of an Aircraft Engine.

PubMed

García, Iker; Przysowa, Radosław; Amorebieta, Josu; Zubia, Joseba

2016-11-11

In this article, we report the design of a reflective intensity-modulated optical fiber sensor for blade tip-clearance measurement, and the experimental results for the first stage of a compressor of an aircraft engine operating in real conditions. The tests were performed in a ground test cell, where the engine completed four cycles from idling state to takeoff and back to idling state. During these tests, the rotational speed of the compressor ranged between 7000 and 15,600 rpm. The main component of the sensor is a tetrafurcated bundle of optical fibers, with which the resulting precision of the experimental measurements was 12 µm for a measurement range from 2 to 4 mm. To get this precision the effect of temperature on the optoelectronic components of the sensor was compensated by calibrating the sensor in a climate chamber. A custom-designed MATLAB program was employed to simulate the behavior of the sensor prior to its manufacture.

Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development and Performance Analysis

NASA Technical Reports Server (NTRS)

Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan; Kirchner, Robert; Engel, Carl D.

2014-01-01

The Space Launch System (SLS) base heating test is broken down into two test programs: (1) Pathfinder and (2) Main Test. The Pathfinder Test Program focuses on the design, development, hot-fire test and performance analyses of the 2% sub-scale SLS core-stage and booster element propulsion systems. The core-stage propulsion system is composed of four gaseous oxygen/hydrogen RS-25D model engines and the booster element is composed of two aluminum-based model solid rocket motors (SRMs). The first section of the paper discusses the motivation and test facility specifications for the test program. The second section briefly investigates the internal flow path of the design. The third section briefly shows the performance of the model RS-25D engines and SRMs for the conducted short duration hot-fire tests. Good agreement is observed based on design prediction analysis and test data. This program is a challenging research and development effort that has not been attempted in 40+ years for a NASA vehicle.

Tip-Clearance Measurement in the First Stage of the Compressor of an Aircraft Engine

PubMed Central

García, Iker; Przysowa, Radosław; Amorebieta, Josu; Zubia, Joseba

2016-01-01

In this article, we report the design of a reflective intensity-modulated optical fiber sensor for blade tip-clearance measurement, and the experimental results for the first stage of a compressor of an aircraft engine operating in real conditions. The tests were performed in a ground test cell, where the engine completed four cycles from idling state to takeoff and back to idling state. During these tests, the rotational speed of the compressor ranged between 7000 and 15,600 rpm. The main component of the sensor is a tetrafurcated bundle of optical fibers, with which the resulting precision of the experimental measurements was 12 µm for a measurement range from 2 to 4 mm. To get this precision the effect of temperature on the optoelectronic components of the sensor was compensated by calibrating the sensor in a climate chamber. A custom-designed MATLAB program was employed to simulate the behavior of the sensor prior to its manufacture. PMID:27845709

Improving of technical characteristics of launch vehicles with liquid rocket engines using active onboard de-orbiting systems

NASA Astrophysics Data System (ADS)

Trushlyakov, V.; Shatrov, Ya.

2017-09-01

In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing. Factual technical characteristics on the example of a promising type of rocket ;Soyuz-2.1.v.; are analyzed. Meeting the TR in relation to anthropogenic impact decrease based on the conventional design approach and the content of the onboard system does not prove to be efficient and leads to depreciation of the initial technical characteristics obtained at the first design stage if these requirements are not included. In this concern, it is shown that the implementation of additional active onboard de-orbiting system (AODS) of worked-off stages (WS) into the onboard LV stages systems allows to meet the TR related to the LV environmental characteristics, including fire-explosion safety. In some cases, the orbital payload mass increases.

Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

NASA Technical Reports Server (NTRS)

Morgan, Morris H., III; Gilinsky, Mikhail M.

2004-01-01

In this project on the first stage (2000-Ol), we continued to develop the previous joint research between the Fluid Mechanics and Acoustics Laboratory (FM&AL) at Hampton University (HU) and the Jet Noise Team (JNT) at the NASA Langley Research Center (NASA LaRC). At the second stage (2001-03), FM&AL team concentrated its efforts on solving of problems of interest to Glenn Research Center (NASA GRC), especially in the field of propulsion system enhancement. The NASA GRC R&D Directorate and LaRC Hyper-X Program specialists in a hypersonic technology jointly with the FM&AL staff conducted research on a wide region of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The last year the Hampton University School of Engineering & Technology was awarded the NASA grant, for creation of the Aeropropulsion Center, and the FM&AL is a key team of the project fulfillment responsible for research in Aeropropulsion and Acoustics (Pillar I). This work is supported by joint research between the NASA GRC/ FM&AL and the Institute of Mechanics at Moscow State University (IMMSU) in Russia under a CRDF grant. The main areas of current scientific interest of the FM&AL include an investigation of the proposed and patented advanced methods for aircraft engine thrust and noise benefits. This is the main subject of our other projects, of which one is presented. The last year we concentrated our efforts to analyze three main problems: (a) new effective methods fuel injection into the flow stream in air-breathing engines; (b) new re-circulation method for mixing, heat transfer and combustion enhancement in propulsion systems and domestic industry application; (c) covexity flow The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines (see, for example, Figures 4). The FM&AL Team uses analytical methods, numerical simulations and experimental tests at the Hampton University campus, NASA and IM/MSU.

Main Chamber Injectors for Advanced Hydrocarbon Booster Engines

NASA Technical Reports Server (NTRS)

Long, Matthew R.; Bazarov, Vladimir G.; Anderson, William E.

2003-01-01

Achieving the highest possible specific impulse has long been a key driver for space launch systems. Recently, more importance has been placed on the need for increased reliability and streamlined launch operations. These general factors along with more specific mission requirements have provided a new focus that is centered on the oxidizer rich staged combustion (ORSC) cycle. Despite a history of use in Russia that extends back to the 1960's, a proven design methodology for ORSC cycle engines does not exist in the West. This lack of design expertise extends to the main chamber injector, a critical subcomponent that largely determines the engine performance and main chamber life. The goals of the effort described here are to establish an empirical knowledge base to provide a fundamental understanding of main chamber injectors and for verification of an injector design methodology for the ORSC cycle. The design of a baseline injector element, derived from information on Russian engines in the open literature, is presented. The baseline injector comprises a gaseous oxidizer core flow and an annular swirling fuel flow. Sets of equations describing the steady-state and the dynamic characteristics of the injector are presented; these equations, which form the basis of the design analysis methodology, will be verified in tests later this year. On-going cold flow studies, using nitrogen and water as simulants, are described which indicate highly atomized and symmetric sprays.

Smart and Intelligent Sensors

NASA Technical Reports Server (NTRS)

Lansaw, John; Schmalzel, John; Figueroa, Jorge

2009-01-01

John C. Stennis Space Center (SSC) provides rocket engine propulsion testing for NASA's space programs. Since the development of the Space Shuttle, every Space Shuttle Main Engine (SSME) has undergone acceptance testing at SSC before going to Kennedy Space Center (KSC) for integration into the Space Shuttle. The SSME is a large cryogenic rocket engine that uses Liquid Hydrogen (LH2) as the fuel. As NASA moves to the new ARES V launch system, the main engines on the new vehicle, as well as the upper stage engine, are currently base lined to be cryogenic rocket engines that will also use LH2. The main rocket engines for the ARES V will be larger than the SSME, while the upper stage engine will be approximately half that size. As a result, significant quantities of hydrogen will be required during the development, testing, and operation of these rocket engines.Better approaches are needed to simplify sensor integration and help reduce life-cycle costs. 1.Smarter sensors. Sensor integration should be a matter of "plug-and-play" making sensors easier to add to a system. Sensors that implement new standards can help address this problem; for example, IEEE STD 1451.4 defines transducer electronic data sheet (TEDS) templates for commonly used sensors such as bridge elements and thermocouples. When a 1451.4 compliant smart sensor is connected to a system that can read the TEDS memory, all information needed to configure the data acquisition system can be uploaded. This reduces the amount of labor required and helps minimize configuration errors. 2.Intelligent sensors. Data received from a sensor be scaled, linearized; and converted to engineering units. Methods to reduce sensor processing overhead at the application node are needed. Smart sensors using low-cost microprocessors with integral data acquisition and communication support offer the means to add these capabilities. Once a processor is embedded, other features can be added; for example, intelligent sensors can make a health assessment to inform the data acquisition client when sensor performance is suspect. 3.Distributed sample synchronization. Networks of sensors require new ways for synchronizing samples. Standards that address the distributed timing problem (for example, IEEE STD 1588) provide the means to aggregate samples from many distributed smart sensors with sub-microsecond accuracy. 4. Reduction in interconnect. Alternative means are needed to reduce the frequent problems associated with cabling and connectors. Wireless technologies offer the promise of reducing interconnects and simultaneously making it easy to quickly add a sensor to a system.

National Aerospace Plane Engine Seals: High Temperature Seal Performance Evaluation

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.

1991-01-01

The key to the successful development of the single stage to orbit National Aerospace Plane (NASP) is the successful development of combined cycle ramjet/scramjet engines that can propel the vehicle to 17,000 mph to reach low Earth orbit. To achieve engine performance over this speed range, movable engine panels are used to tailor engine flow that require low leakage, high temperature seals around their perimeter. NASA-Lewis is developing a family of new high temperature seals to form effective barriers against leakage of extremely hot (greater than 2000 F), high pressure (up to 100 psi) flow path gases containing hydrogen and oxygen. Preventing backside leakage of these explosive gas mixtures is paramount in preventing the potential loss of the engine or the entire vehicle. Seal technology development accomplishments are described in the three main areas of concept development, test, and evaluation and analytical development.

The J-2X Fuel Turbopump - Design, Development, and Test

NASA Technical Reports Server (NTRS)

Tellier, James G.; Hawkins, Lakiesha V.; Shinguchi, Brian H.; Marsh, Matthew W.

2011-01-01

Pratt and Whitney Rocketdyne (PWR), a NASA subcontractor, is executing the design, development, test, and evaluation (DDT&E) of a liquid oxygen, liquid hydrogen two hundred ninety four thousand pound thrust rocket engine initially intended for the Upper Stage (US) and Earth Departure Stage (EDS) of the Constellation Program Ares-I Crew Launch Vehicle (CLV). A key element of the design approach was to base the new J-2X engine on the heritage J-2S engine with the intent of uprating the engine and incorporating SSME and RS-68 lessons learned. The J-2S engine was a design upgrade of the flight proven J-2 configuration used to put American astronauts on the moon. The J-2S Fuel Turbopump (FTP) was the first Rocketdyne-designed liquid hydrogen centrifugal pump and provided many of the early lessons learned for the Space Shuttle Main Engine High Pressure Fuel Turbopumps. This paper will discuss the design trades and analyses performed for the current J-2X FTP to increase turbine life; increase structural margins, facilitate component fabrication; expedite turbopump assembly; and increase rotordynamic stability margins. Risk mitigation tests including inducer water tests, whirligig turbine blade tests, turbine air rig tests, and workhorse gas generator tests characterized operating environments, drove design modifications, or identified performance impact. Engineering design, fabrication, analysis, and assembly activities support FTP readiness for the first J-2X engine test scheduled for July 2011.

Dynamic characteristics of single crystal SSME blades

NASA Technical Reports Server (NTRS)

Moss, L. A.; Smith, T. E.

1987-01-01

The Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) blades are currently manufactured using a directionally solidified (DS) material, MAR-M-246+Hf. However, a necessity to reduce the occurrence of fatigue cracking within the DS blades has lead to an interest in the use of a single crystal (SC) material, PWA-1480. A study was initiated to determine the dynamic characteristics of the HPFTP blades made of SC material and find possible critical engine order excitations. This study examined both the first and second stage drive turbine blades of the HPFTP. The dynamic characterization was done analytically as well as experimentally. The analytical study examined the SC first stage HPFTP blade dynamic characteristics under typical operating conditions. The blades were analyzed using MSC/NASTRAN and a finite element model. Two operating conditions, 27500 RPM and 35000 RPM, were investigated.

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 7: High pressure fuel turbo-pump third stage impeller analysis

NASA Technical Reports Server (NTRS)

Pool, Kirby V.

1989-01-01

This volume summarizes the analysis used to assess the structural life of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbo-Pump (HPFTP) Third Stage Impeller. This analysis was performed in three phases, all using the DIAL finite element code. The first phase was a static stress analysis to determine the mean (non-varying) stress and static margin of safety for the part. The loads involved were steady state pressure and centrifugal force due to spinning. The second phase of the analysis was a modal survey to determine the vibrational modes and natural frequencies of the impeller. The third phase was a dynamic response analysis to determine the alternating component of the stress due to time varying pressure impulses at the outlet (diffuser) side of the impeller. The results of the three phases of the analysis show that the Third Stage Impeller operates very near the upper limits of its capability at full power level (FPL) loading. The static loading alone creates stresses in some areas of the shroud which exceed the yield point of the material. Additional cyclic loading due to the dynamic force could lead to a significant reduction in the life of this part. The cyclic stresses determined in the dynamic response phase of this study are based on an assumption regarding the magnitude of the forcing function.

Heat transfer and pressure measurements for the SSME fuel turbine

NASA Technical Reports Server (NTRS)

Dunn, Michael G.; Kim, Jungho

1991-01-01

A measurement program is underway using the Rocketdyne two-stage Space Shuttle Main Engine (SSME) fuel turbine. The measurements use a very large shock tunnel to produce a short-duration source of heated and pressurized gas which is subsequently passed through the turbine. Within this environment, the turbine is operated at the design values of flow function, stage pressure ratio, stage temperature ratio, and corrected speed. The first stage vane row and the first stage blade row are instrumented in both the spanwise and chordwise directions with pressure transducers and heat flux gages. The specific measurements to be taken include time averaged surface pressure and heat flux distributions on the vane and blade, flow passage static pressure, flow passage total pressure and total temperature distributions, and phase resolved surface pressure and heat flux on the blade.

Educational-research laboratory "electric circuits" on the base of digital technologies

NASA Astrophysics Data System (ADS)

Koroteyev, V. I.; Florentsev, V. V.; Florentseva, N. I.

2017-01-01

The problem of research activity of trainees' activation in the educational-research laboratory "Electric Circuits" using innovative methodological solutions and digital technologies is considered. The main task is in creation of the unified experimental research information-educational environment "Electrical Engineering". The problems arising during the developing and application of the modern software and hardware, experimental and research stands and digital control and measuring systems are presented. This paper presents the main stages of development and creation of educational-research laboratory "Electrical Circuits" at the Department of Electrical Engineering of NRNU MEPhI. The authors also consider the analogues of the described research complex offered by various educational institutions and companies. The analysis of their strengths and weaknesses, on which the advantages of the proposed solution are based, is held.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 Engine major components with callouts (chart 1): The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 engine major components with callouts (chart 1). The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Space Shuttle Main Engine - The Relentless Pursuit of Improvement

NASA Technical Reports Server (NTRS)

VanHooser, Katherine P.; Bradley, Douglas P.

2011-01-01

The Space Shuttle Main Engine (SSME) is the only reusable large liquid rocket engine ever developed. The specific impulse delivered by the staged combustion cycle, substantially higher than previous rocket engines, minimized volume and weight for the integrated vehicle. The dual pre-burner configuration permitted precise mixture ratio and thrust control while the fully redundant controller and avionics provided a very high degree of system reliability and health diagnosis. The main engine controller design was the first rocket engine application to incorporate digital processing. The engine was required to operate at a high chamber pressure to minimize engine volume and weight. Power level throttling was required to minimize structural loads on the vehicle early in flight and acceleration levels on the crew late in ascent. Fatigue capability, strength, ease of assembly and disassembly, inspectability, and materials compatibility were all major considerations in achieving a fully reusable design. During the multi-decade program the design evolved substantially using a series of block upgrades. A number of materials and manufacturing challenges were encountered throughout SSME s history. Significant development was required for the final configuration of the high pressure turbopumps. Fracture control was implemented to assess life limits of critical materials and components. Survival in the hydrogen environment required assessment of hydrogen embrittlement. Instrumentation systems were a challenge due to the harsh thermal and dynamic environments within the engine. Extensive inspection procedures were developed to assess the engine components between flights. The Space Shuttle Main Engine achieved a remarkable flight performance record. All flights were successful with only one mission requiring an ascent abort condition, which still resulted in an acceptable orbit and mission. This was achieved in large part via extensive ground testing to fully characterize performance and to establish acceptable life limits. During the program over a million seconds of accumulated test and flight time was achieved. Post flight inspection and assessment was a key part of assuring proper performance of the flight hardware. By the end of the program the predicted reliability had improved by a factor of four. These unique challenges, evolution of the design, and the resulting reliability will be discussed in this paper.

SSME single-crystal turbine blade dynamics

NASA Technical Reports Server (NTRS)

Moss, Larry A.

1988-01-01

A study was performrd to determine the dynamic characteristics of the Space Shuttle Main Engine high pressure fuel turbopump (HPFTP) blades made of single crystal (SC) material. The first and second stage drive turbine blades of HPFTP were examined. The nonrotating natural frequencies were determined experimentally and analytically. The experimental results of the SC second stage blade were used to verify the analytical procedures. The study examined the SC first stage blade natural frequencies with respect to crystal orientation at typical operating conditions. The SC blade dynamic response was predicted to be less than the directionally solidified base. Crystal axis orientation optimization indicated that the third mode interference will exist in any SC orientation.

Laser velocimeter measurements of the flow downstream of the Space Shuttle Main Engine high pressure oxidizer turbopump first-stage turbine nozzle

NASA Technical Reports Server (NTRS)

Ferguson, T. V.; Havskjold, G. L.; Rojas, L.

1988-01-01

A laser two-focus velocimeter was used in an open-loop water test facility in order to map the flowfield downstream of the SSME's high-pressure oxidizer turbopump first-stage turbine nozzle; attention was given to the effects of the upstream strut-downstream nozzle configuration on the flow at the rotor inlet, in order to estimate dynamic loads on the first-stage rotor blades. Velocity and flow angles were plotted as a function of circumferential position, and were found to clearly display the periodic behavior of the wake flow field. The influence of the upstream centerbody-supporting struts on the vane nozzle wake pattern was evident.

SSME single crystal turbine blade dynamics

NASA Technical Reports Server (NTRS)

Moss, Larry A.; Smith, Todd E.

1987-01-01

A study was performed to determine the dynamic characteristics of the Space Shuttle main engine high pressure fuel turbopump (HPFTP) blades made of single crystal (SC) material. The first and second stage drive turbine blades of HPFTP were examined. The nonrotating natural frequencies were determined experimentally and analytically. The experimental results of the SC second stage blade were used to verify the analytical procedures. The analytical study examined the SC first stage blade natural frequencies with respect to crystal orientation at typical operating conditions. The SC blade dynamic response was predicted to be less than the directionally solidified blade. Crystal axis orientation optimization indicated the third mode interference will exist in any SC orientation.

A Framework for Assessing the Reusability of Hardware (Reusable Rocket Engines)

NASA Technical Reports Server (NTRS)

Childress-Thompson, Rhonda; Thomas, Dale; Farrington, Philip

2016-01-01

Within the past few years, there has been a renewed interest in reusability as it applies to space flight hardware. Commercial companies such as Space Exploration Technologies Corporation (SpaceX), Blue Origin, and United Launch Alliance (ULA) are pursuing reusable hardware. Even foreign companies are pursuing this option. The Indian Space Research Organization (ISRO) launched a reusable space plane technology demonstrator and Airbus Defense and Space is planning to recover the main engines and avionics from its Advanced Expendable Launcher with Innovative engine Economy [1] [2]. To date, the Space Shuttle remains as the only Reusable Launch (RLV) to have flown repeated missions and the Space Shutte Main Engine (SSME) is the only demonstrated reusable engine. Whether the hardware being considered for reuse is a launch vehicle (fully reusable), a first stage (partially reusable), or a booster engine (single component), the overall governing process is the same; it must be recovered and recertified for flight. Therefore, there is a need to identify the key factors in determining the reusability of flight hardware. This paper begins with defining reusability to set the context, addresses the significance of reuse, and discusses areas that limit successful implementation. Finally, this research identifies the factors that should be considered when incorporating reuse.

Computational Plume Modeling of COnceptual ARES Vehicle Stage Tests

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Ahuja, Vineet

2007-01-01

The plume-induced environment of a conceptual ARES V vehicle stage test at the NASA Stennis Space Center (NASA-SSC) was modeled using computational fluid dynamics (CFD). A full-scale multi-element grid was generated for the NASA-SSC B-2 test stand with the ARES V stage being located in a proposed off-center forward position. The plume produced by the ARES V main power plant (cluster of five RS-68 LOX/LH2 engines) was simulated using a multi-element flow solver - CRUNCH. The primary objective of this work was to obtain a fundamental understanding of the ARES V plume and its impingement characteristics on the B-2 flame-deflector. The location, size and shape of the impingement region were quantified along with the un-cooled deflector wall pressures, temperatures and incident heating rates. Issues with the proposed tests were identified and several of these addressed using the CFD methodology. The final results of this modeling effort will provide useful data and boundary conditions in upcoming engineering studies that are directed towards determining the required facility modifications for ensuring safe and reliable stage testing in support of the Constellation Program.

Fluid-Structure Interaction Modeling of the Reefed Stages of the Orion Spacecraft Main Parachutes

NASA Astrophysics Data System (ADS)

Boswell, Cody W.

Spacecraft parachutes are typically used in multiple stages, starting with a "reefed" stage where a cable along the parachute skirt constrains the diameter to be less than the diameter in the subsequent stage. After a certain period of time during the descent, the cable is cut and the parachute "disreefs" (i.e. expands) to the next stage. Computing the parachute shape at the reefed stage and fluid-structure interaction (FSI) modeling during the disreefing involve computational challenges beyond those we have in FSI modeling of fully-open spacecraft parachutes. These additional challenges are created by the increased geometric complexities and by the rapid changes in the parachute geometry. The computational challenges are further increased because of the added geometric porosity of the latest design, where the "windows" created by the removal of panels and the wider gaps created by the removal of sails compound the geometric and flow complexity. Orion spacecraft main parachutes will have three stages, with computation of the Stage 1 shape and FSI modeling of disreefing from Stage 1 to Stage 2 being the most challenging. We present the special modeling techniques we devised to address the computational challenges and the results from the computations carried out. We also present the methods we devised to calculate for a parachute gore the radius of curvature in the circumferential direction. The curvature values are intended for quick and simple engineering analysis in estimating the structural stresses.

Advanced space engine preliminary design

NASA Technical Reports Server (NTRS)

Cuffe, J. P. B.; Bradie, R. E.

1973-01-01

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

Milestone report TCTP application to the SSME hydrogen system analysis

NASA Technical Reports Server (NTRS)

Richards, J. S.

1975-01-01

The Transient Cryogen Transfer Computer Program (TCTP) developed and verified for LOX systems by analyses of Skylab S-1B stage loading data from John F. Kennedy Space Center launches was extended to include hydrogen as the working fluid. The feasibility of incorporating TCTP into the space shuttle main engine dynamic model was studied. The program applications are documented.

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/Hydrogen Rocket Engine

NASA Technical Reports Server (NTRS)

Sutton, R. F.; Lariviere, B. W.

1993-01-01

An advanced LOX/LH2 expander cycle rocket engine, producing 15,000 lbf thrust for Orbital Transfer Vehicle missions, was tested to determine ignition, transition, and main stage characteristics. Detail design and fabrication of the pump fed RS44 integrated component evaluator (ICE) was accomplished using company discretionary resources and was tested under this contracted effort. Successful demonstrations were completed to about the 50 percent fuel turbopump power level (87,000 RPM), but during this last test, a high pressure fuel turbopump (HPFTP) bearing failed curtailing the test program. No other hardware were affected by the HPFTP premature shutdown. The ICE operations matched well with the predicted start transient simulations. The tests demonstrated the feasibility of a high performance advanced expander cycle engine. All engine components operated nominally, except for the HPFTP, during the engine hot-fire tests. A failure investigation was completed using company discretionary resources.

Access to space

NASA Astrophysics Data System (ADS)

1994-07-01

The goal of this conceptual design was to devise a reusable, commercially viable, single-stage-to-orbit vehicle. The vehicle has the ability to deliver a 9100 kg (20,000 lb) payload to a low earth orbit of 433 km to 933 km (250 n.mi. - 450 n.mi.). The SSTO vehicle is 51 meters in length and has a gross takeoff mass of 680,400 kg (1,500,000 lb). The vehicle incorporates three RD-701 engines for the main propulsion system and two RL-10 engines for the orbital maneuvering system. The vehicle is designed for a three day stay on orbit with two crew members.

Nonlinear heat transfer and structural analyses of SSME turbine blades

NASA Technical Reports Server (NTRS)

Abdul-Aziz, A.; Kaufman, A.

1987-01-01

Three-dimensional nonlinear finite-element heat transfer and structural analyses were performed for the first stage high-pressure fuel turbopump blade of the space shuttle main engine (SSME). Directionally solidified (DS) MAR-M 246 material properties were considered for the analyses. Analytical conditions were based on a typical test stand engine cycle. Blade temperature and stress-strain histories were calculated using MARC finite-element computer code. The study was undertaken to assess the structural response of an SSME turbine blade and to gain greater understanding of blade damage mechanisms, convective cooling effects, and the thermal-mechanical effects.

Access to space

NASA Technical Reports Server (NTRS)

1994-01-01

The goal of this conceptual design was to devise a reusable, commercially viable, single-stage-to-orbit vehicle. The vehicle has the ability to deliver a 9100 kg (20,000 lb) payload to a low earth orbit of 433 km to 933 km (250 n.mi. - 450 n.mi.). The SSTO vehicle is 51 meters in length and has a gross takeoff mass of 680,400 kg (1,500,000 lb). The vehicle incorporates three RD-701 engines for the main propulsion system and two RL-10 engines for the orbital maneuvering system. The vehicle is designed for a three day stay on orbit with two crew members.

50th anniversary logo

NASA Image and Video Library

2011-03-02

John C. Stennis Space Center is celebrating its 50th anniversary in 2011. NASA announced plans to build a rocket engine test facility in Hancock County, Miss., on Oct. 25, 1961. A new anniversary logo highlights the theme of the anniversary year - celebrating Stennis as a unique federal city and its five decades of powering America's space dreams. Stennis is home to more than 30 federal, state, academic and private organizations and several technology-based companies. In addition to testing Apollo Program rocket stages that carried humans to the moon, Stennis tested every main engine used in more than 130 space shuttle flights.

A Rainbow View of NASA's RS-25 Engine Test

NASA Image and Video Library

2017-02-22

NASA engineers conducted their first RS-25 test of 2017 on the A-1 Test Stand at Stennis Space Center near Bay St. Louis, Mississippi, on Feb. 22, continuing to collect data on the performance of the rocket engine that will help power the new Space Launch System (SLS) rocket. Shown from the viewpoint of an overhead drone, the test of development engine No. 0528 ran the scheduled 380 seconds (six minutes and 20 seconds), allowing engineers to monitor various engine operating conditions. The test represents another step forward in development of the rocket that will launch humans aboard Orion deeper into space than ever before. Four RS-25 engines, together with a pair of solid rocket boosters, will power the SLS at launch on its deep-space missions. The engines for the first four SLS flights are former space shuttle main engines, which were tested extensively at Stennis and are some of the most proven engines in the world. Engineers are conducting an ongoing series of tests this year for SLS on both development and flight engines for future flights to ensure the engine, outfitted with a new controller, can perform at the higher level under a variety of conditions and situations. Stennis is also preparing its B-2 Test Stand to test the core stage for the first SLS flight with Orion, known as Exploration Mission-1. That testing will involve installing the flight stage on the stand and firing its four RS-25 engines simultaneously, just as during an actual launch. The Feb. 22 test was conducted by Aerojet Rocketdyne and Syncom Space Services engineers and operators. Aerojet Rocketdyne is the prime contractor for the RS-25 engines. Syncom Space Services is the prime contractor for Stennis facilities and operations. PAO Name:Kim Henry Phone Number:256-544-1899 Email Address: kimberly.m.henry@nasa.gov

Analysis of space tug operating techniques. Volume 2: Study results

NASA Technical Reports Server (NTRS)

1972-01-01

The design requirements for space tug systems and cost analysis of the refurbishment phases are discussed. The vehicle is an integral propulsion stage using liquid hydrogen and liquid oxygen as propellants and is capable of operating either as a fully or a partially autonomous vehicle. Structural features are an integral liquid hydrogen tank, a liquid oxygen tank, a meteoroid shield, an aft conical docking and structural support ring, and a staged combustion main engine. The vehicle is constructed of major modules for ease of maintenance. Line drawings and block diagrams are included to explain the maintenance requirements for the subsystems.

Space shuttle vehicle rocket plume impingement study for separation analysis. Tasks 2 and 3: Definition and preliminary plume impingement analysis for the MSC booster

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.; Penny, M. M.; Prozan, R. J.

1970-01-01

The results are presented of a space shuttle plume impingement study for the Manned Spacecraft Center configuration. This study was conducted as two tasks which were to (1) define the orbiter main stage engine exhaust plume flow field, and (2) define the plume impingement heating, force and resulting moment environments on the booster during the staging maneuver. To adequately define these environments during the staging maneuver and allow for deviation from the nominal separation trajectory, a multitude of relative orbiter/booster positions are analyzed which map the region that contains the separation trajectories. The data presented can be used to determine a separation trajectory which will result in acceptable impingement heating rates, forces, and the resulting moments. The data, presented in graphical form, include the effect of roll, pitch and yaw maneuvers for the booster. Quasi-steady state analysis methods were used with the orbiter engine operating at full thrust. To obtain partial thrust results, simple ratio equations are presented.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. The H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. Each H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

The use of participant-observation protocol in an industrial engineering research.

PubMed

Silveira e Silva, Renato da; Sznelwar, Laerte Idal; D'Afonseca e Silva, Victor

2012-01-01

Based on literature, this article aims to present the "participant-observation' research protocol, and its practical application in the industrial engineering field, more specifically within the area of design development, and in the case shown by this article, of interiors' design. The main target is to identify the concept of the method, i.e., from its characteristics to structure a general sense about the subject, so that the protocol can be used in different areas of knowledge, especially those ones which are committed with the scientific research involving the expertise from researchers, and subjective feelings and opinions of the users of an engineering product, and how this knowledge can be benefic for product design, contributing since the earliest stage of design.

Saturn Apollo Program

NASA Image and Video Library

1965-03-04

Pictured is a J-2 engine being processed at Marshall Space Flight Center (MSFC). A single J-2 engine was utilized on the S-IVB stage, the second stage of the Saturn IB and the third stage of the Saturn V vehicles, while a cluster of five J-2 engines powered the second (S-II) stage of the Saturn V launch vehicle. The Saturn V was designed, developed, and tested by engineers at MSFC.

Theoretical study of short pile effect in tunnel excavation

NASA Astrophysics Data System (ADS)

Tian, Xiao-yan; Liu, Jing; Gao, Xiao-mei; Li, Yuan

2017-09-01

The Misaki Sato Go ideal elastoplastic model is adopted and the two stage analysis theory is used to study the effect of tunnel excavation on short pile effect in this paper. In the first stage, the free field vertical displacement of the soil at the corresponding pile location is obtained by using empirical formula. In the second stage, the displacement is applied to the corresponding pile location. The equilibrium condition of micro physical differential equation settlement of piles. Then through logical deduction and the boundary condition expressions of the settlement calculation, obtain the pile side friction resistance and axial force of the week. Finally, an engineering example is used to analyze the influence of the change of main parameters on their effects.

Pollution reduction technology program for class T4(JT8D) engines

NASA Technical Reports Server (NTRS)

Roberts, R.; Fiorentino, A. J.; Diehl, L. A.

1977-01-01

The technology required to develop commercial gas turbine engines with reduced exhaust emissions was demonstrated. Can-annular combustor systems for the JT8D engine family (EPA class T4) were investigated. The JT8D turbofan engine is an axial-flow, dual-spool, moderate-bypass-ratio design. It has a two-stage fan, a four-stage low-pressure compressor driven by a three-stage low-pressure turbine, and a seven-stage high-pressure compressor driven by a single-stage high-pressure turbine. A cross section of the JT8D-17 showing the mechanical configuration is given. Key specifications for this engine are listed.

Rocketdyne - J-2 Saturn V 2nd and 3rd Stage Engine. Chapter 2, Appendix D

NASA Technical Reports Server (NTRS)

Coffman, Paul

2009-01-01

The J-2 engine was unique in many respects. Technology was not nearly as well-developed in oxygen/hydrogen engines at the start of the J-2 project. As a result, it experienced a number of "teething" problems. It was used in two stages on the Saturn V vehicle in the Apollo Program, as well as on the later Skylab and Apollo/Soyuz programs. In the Apollo Program, it was used on the S-II stage, which was the second stage of the Saturn V vehicle. There were five J-2 engines at the back end of the S-II Stage. In the S-IV-B stage, it was a single engine, but that single engine had to restart. The Apollo mission called for the entire vehicle to reach orbital velocity in low Earth orbit after the first firing of the Saturn-IV-B stage and, subsequently, to fire a second time to go on to the moon. The engine had to be man-rated (worthy of transporting humans). It had to have a high thrust rate and performance associated with oxygen/hydrogen engines, although there were some compromises there. It had to gimbal for thrust vector control. It was an open-cycle gas generator engine delivering up to 230,000 pounds of thrust.

Propulsion Risk Reduction Activities for Non-Toxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for non-toxic or "green" propellants. The PCAD project focuses on the development of non-toxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of non-toxic propellants for space missions. Implementation of non-toxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that non-toxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Propulsion Risk Reduction Activities for Nontoxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth L.

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for nontoxic or "green" propellants. The PCAD project focuses on the development of nontoxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of nontoxic propellants for space missions. Implementation of nontoxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Space Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1975-07-21

Ken Baskin, an engineer from the Facilities and Engineering Branch at the National Aeronautics and Space Administration’s (NASA) Lewis Research Center checks a complete 2.25-scale model of the shuttle in the 10- by 10-Foot Supersonic Wind Tunnel. Baskin’s space shuttle project began in July 1976 during the run-up to the shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Rockwell International engineers needed to understand these issues in order to design proper thermal protection. The 10- by 10 tests evaluated the base heating and pressure. The test’s specific objectives were to measure heat transfer and pressure distributions around the orbiter’s external tank and solid rocket booster afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions due to rocket exhaust-induced flow separation, and determine gas recovery temperatures using gas temperature probes and heated model base components. The shuttle model’s main engines and solid rockets were fired during the tests, then just the main engines in an effort to simulate a launch. The researchers conducted 163 runs in the 10- by 10 during the test program.

Apollo-Lunar Orbital Rendezvous Technique

NASA Technical Reports Server (NTRS)

1963-01-01

The film shows artists rendition of the spacecrafts, boosters, and flight of the Apollo lunar missions. The Apollo spacecraft will consist of three modules: the manned Command Module; the Service Module, which contains propulsion systems; and the Lunar Excursion Module (LEM) to carry astronauts to the moon and back to the Command and Service Modules. The spacecraft will be launched via a three-stage Saturn booster. The first stage will provide 7.5 million pounds of thrust from five F-1 engines for liftoff and initial powered flight. The second stage will develop 1 million pounds of thrust from five J-2 engines to boost the spacecraft almost into Earth orbit. Immediately after ignition of the second stage, the Launch Escape System will be jettisoned. A single J-2 engine in the S4B stage will provide 200,000 pounds of thrust to place the spacecraft in an earth parking orbit. It also will be used to propel the spacecraft into a translunar trajectory, then it will separate from the Apollo Modules. Onboard propulsion systems will be used to insert the spacecraft into lunar orbit. Two astronauts will enter the LEM, which will separate from the command and service modules. The LEM will go into elliptical orbit and prepare for landing. The LEM will lift off of the Moon's surface to return to the Command and Service Modules, and most likely be left in lunar orbit. After leaving the Moon's orbit, and shortly before entering Earth's orbit, the Service Module will be ejected. The Command Module will be oriented for reentry into the Earth's atmosphere. A drogue parachute will deploy at approximately 50,000 feet, followed by the main parachute system for touchdown.

Two-stage earth-to-orbit vehicles with dual-fuel propulsion in the Orbiter

NASA Technical Reports Server (NTRS)

Martin, J. A.

1982-01-01

Earth-to-orbit vehicle studies of future replacements for the Space Shuttle are needed to guide technology development. Previous studies that have examined single-stage vehicles have shown advantages for dual-fuel propulsion. Previous two-stage system studies have assumed all-hydrogen fuel for the Orbiters. The present study examined dual-fuel Orbiters and found that the system dry mass could be reduced with this concept. The possibility of staging the booster at a staging velocity low enough to allow coast-back to the launch site is shown to be beneficial, particularly in combination with a dual-fuel Orbiter. An engine evaluation indicated the same ranking of engines as did a previous single-stage study. Propane and RP-1 fuels result in lower vehicle dry mass than methane, and staged-combustion engines are preferred over gas-generator engines. The sensitivity to the engine selection is less for two-stage systems than for single-stage systems.

Large liquid rocket engine transient performance simulation system

NASA Technical Reports Server (NTRS)

Mason, J. R.; Southwick, R. D.

1991-01-01

A simulation system, ROCETS, was designed and developed to allow cost-effective computer predictions of liquid rocket engine transient performance. The system allows a user to generate a simulation of any rocket engine configuration using component modules stored in a library through high-level input commands. The system library currently contains 24 component modules, 57 sub-modules and maps, and 33 system routines and utilities. FORTRAN models from other sources can be operated in the system upon inclusion of interface information on comment cards. Operation of the simulation is simplified for the user by run, execution, and output processors. The simulation system makes available steady-state trim balance, transient operation, and linear partial generation. The system utilizes a modern equation solver for efficient operation of the simulations. Transient integration methods include integral and differential forms for the trapezoidal, first order Gear, and second order Gear corrector equations. A detailed technology test bed engine (TTBE) model was generated to be used as the acceptance test of the simulation system. The general level of model detail was that reflected in the Space Shuttle Main Engine DTM. The model successfully obtained steady-state balance in main stage operation and simulated throttle transients, including engine starts and shutdown. A NASA FORTRAN control model was obtained, ROCETS interface installed in comment cards, and operated with the TTBE model in closed-loop transient mode.

Composite engines for application to a single-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

Bendot, J. G.; Brown, P. N.; Piercy, T. G.

1975-01-01

Seven composite engines were designed for application to a reusable single-stage-to-orbit vehicle. The engine designs were variations of the supercharged ejector ramjet engine. The resulting performance, weight, and drawings of each engine form a data base for establishing a potential of this class of composite engine to various missions, including the single-stage-to-orbit application. The impact of advanced technology in the design of the critical fan turbine was established.

Remembering the Giants: Apollo Rocket Propulsion Development

NASA Technical Reports Server (NTRS)

Fisher, Steven C. (Editor); Rahman, Shamim A. (Editor)

2009-01-01

Topics discussed include: Rocketdyne - F-1 Saturn V First Stage Engine; Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine; Rocketdyne - SE-7 & SE-8 Engines; Aerojet - AJ10-137 Apollo Service Module Engine; Aerojet - Attitude Control Engines; TRW - Lunar Descent Engine; and Rocketdyne - Lunar Ascent Engine.

Physics based modeling of axial compressor stall

NASA Astrophysics Data System (ADS)

Zaki, Mina Adel

2009-12-01

Axial compressors are used in a wide variety of aerodynamic applications and are one of the most important components in aero-engines. However, the operability of compressors is limited at low-mass flow rates by fluid dynamic instabilities such as stall and surge. These instabilities can lead to engine failure and loss of engine power which can compromise the aircraft safety and reliability. Thus, a better understanding of how stall occurs and the causes behind its inception is extremely important. In the vicinity of the stall line, the flow field is inherently unsteady due to the interactions between adjacent rows of blades, formation of separation cells, and the viscous effects including shock-boundary layer interactions. Accurate modeling of these phenomena requires a proper set of stable and accurate boundary conditions at the rotor-stator interface that conserve mass, momentum, and energy, while eliminating false reflections. As a part of this research effort, an existing 3-D Navier-Stokes analysis for modeling single stage compressors has been modified to model multi-stage axial compressors and turbines. Several rotor-stator interface boundary conditions have been implemented. These conditions have been evaluated for the first stage (a stator and a rotor) of the two-stage fuel turbine on the space shuttle main engine (SSME). Their effectiveness in conserving global properties such as mass, momentum, and energy across the interface while yielding good performance predictions has been evaluated. While all the methods gave satisfactory results, a characteristic based approach and an unsteady sliding mesh approach are found to work best. Accurate modeling of the formation of stall cells requires the use of advanced turbulence models. As a part of this effort, a new advanced turbulence model called the Hybrid RANS/KES (HRKES) model has been developed and implemented. This model solves the Menter's k-o-SST model near walls and switches to the Kinetic Eddy Simulation (KES) model away from walls. The KES model solves directly for local turbulent kinetic energy and local turbulent length scales, alleviating the grid spacing dependency of the length scales found in other Detached Eddy Simulation (DES) and Hybrid RANS/LES (HRLES) models. Within the HRKES model, combinations of two different blending functions have been evaluated for integrating the near wall model with the KES model. The use of realizability constraints to bound the KES model parameters has also been studied for several internal and external flows. The current methodology is used in the prediction of the performance map for the NASA Stage 35 compressor configuration as a representative of a modern compressor stage. The present approach is found to effectively predict the onset of stall. It is found that the rotor blade tip leakage vortex and its interaction with the shock wave is mainly the reason behind the stall inception in this compressor stage.

High-End Concept Based on Hypersonic Two-Stage Rocket and Electro-Magnetic Railgun to Launch Micro-Satellites Into Low-Earth

NASA Astrophysics Data System (ADS)

Bozic, O.; Longo, J. M.; Giese, P.; Behren, J.

2005-02-01

The electromagnetic railgun technology appears to be an interesting alternative to launch small payloads into Low Earth Orbit (LEO), as this may introduce lower launch costs. A high-end solution, based upon present state of the art technology, has been investigated to derive the technical boundary conditions for the application of such a new system. This paper presents the main concept and the design aspects of such propelled projectile with special emphasis on flight mechanics, aero-/thermodynamics, materials and propulsion characteristics. Launch angles and trajectory optimisation analyses are carried out by means of 3 degree of freedom simulations (3DOF). The aerodynamic form of the projectile is optimised to provoke minimum drag and low heat loads. The surface temperature distribution for critical zones is calculated with DLR developed Navier-Stokes codes TAU, HOTSOSE, whereas the engineering tool HF3T is used for time dependent calculations of heat loads and temperatures on project surface and inner structures. Furthermore, competing propulsions systems are considered for the rocket engines of both stages. The structural mass is analysed mostly on the basis of carbon fibre reinforced materials as well as classical aerospace metallic materials. Finally, this paper gives a critical overview of the technical feasibility and cost of small rockets for such missions. Key words: micro-satellite, two-stage-rocket, railgun, rocket-engines, aero/thermodynamic, mass optimization

Lunar lander and return propulsion system trade study

NASA Technical Reports Server (NTRS)

Hurlbert, Eric A.; Moreland, Robert; Sanders, Gerald B.; Robertson, Edward A.; Amidei, David; Mulholland, John

1993-01-01

This trade study was initiated at NASA/JSC in May 1992 to develop and evaluate main propulsion system alternatives to the reference First Lunar Outpost (FLO) lander and return-stage transportation system concept. Thirteen alternative configurations were developed to explore the impacts of various combinations of return stage propellants, using either pressure or pump-fed propulsion systems and various staging options. Besides two-stage vehicle concepts, the merits of single-stage and stage-and-a-half options were also assessed in combination with high-performance liquid oxygen and liquid hydrogen propellants. Configurations using an integrated modular cryogenic engine were developed to assess potential improvements in packaging efficiency, mass performance, and system reliability compared to non-modular cryogenic designs. The selection process to evaluate the various designs was the analytic hierarchy process. The trade study showed that a pressure-fed MMH/N2O4 return stage and RL10-based lander stage is the best option for a 1999 launch. While results of this study are tailored to FLO needs, the design date, criteria, and selection methodology are applicable to the design of other crewed lunar landing and return vehicles.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This image illustrates the basic differences between the three Saturn launch vehicles developed by the Marshall Space Flight Center. The Saturn I, consisted of two stages, the S-I (eight H-1 engines) and the S-IV (six RL-10 engines). The Saturn IB (center) also consisted of two stages, the S-IB (eight H-1 engines) and the S-IVB (one J-2 engine). The Saturn V consisted of three stages, the S-IC (five F-1 engines), the S-II (five J-2 engines), and the S-IVB (one J-2 engine).

Evaluation of Proposed Rocket Engines for Earth-to-Orbit Vehicles

NASA Technical Reports Server (NTRS)

Martin, James A.; Kramer, Richard D.

1990-01-01

The objective is to evaluate recently analyzed rocket engines for advanced Earth-to-orbit vehicles. The engines evaluated are full-flow staged combustion engines and split expander engines, both at mixture ratios at 6 and above with oxygen and hydrogen propellants. The vehicles considered are single-stage and two-stage fully reusable vehicles and the Space Shuttle with liquid rocket boosters. The results indicate that the split expander engine at a mixture ratio of about 7 is competitive with the full-flow staged combustion engine for all three vehicle concepts. A key factor in this result is the capability to increase the chamber pressure for the split expander as the mixture ratio is increased from 6 to 7.

Perfusion properties of scaffolds: A new approach to tissue engineering designs for bone regeneration

NASA Astrophysics Data System (ADS)

Larionov, P. M.; Maslov, N. A.; Papaeva, E. O.; Yunoshev, A. S.; Filipenko, M. L.; Bogachev, S. S.; Proskurina, A. S.; Samokhin, A. G.; Kudrov, G. A.; Tereshchenko, V. P.; Pavlov, V. V.; Mihailovsky, M. V.; Prohorenko, V. M.; Titov, A. T.; Mamonova, E. V.; Sadovoy, M. A.

2017-09-01

The main approach to tissue engineering involves the use of scaffolds seeded with cells, followed by culturing in a bioreactor. However, the effective use of a bioreactor requires adaptation of the scaffold at the stage of its design. In our opinion, this means assessment of the perfusion properties of the scaffold. Transverse and longitudinal perfusion under hydrostatic pressure of 5, 10, and 15 mmHg, as well as the significance of electrospinning parameters for fabrication of a scaffold sheet and the composition of composite material—11% w/v polycaprolactone with gelatinization of 0.5%, 2%, and 4%, were demonstrated.

Advanced transportation system studies technical area 3: Alternate propulsion subsystem concepts, volume 2

NASA Technical Reports Server (NTRS)

Levak, Daniel

1993-01-01

The Alternate Propulsion Subsystem Concepts contract had five tasks defined for the first year. The tasks were: F-1A Restart Study, J-2S Restart Study, Propulsion Database Development, Space Shuttle Main Engine (SSME) Upper Stage Use, and CER's for Liquid Propellant Rocket Engines. The detailed study results, with the data to support the conclusions from various analyses, are being reported as a series of five separate Final Task Reports. Consequently, this volume only reports the required programmatic information concerning Computer Aided Design Documentation, and New Technology Reports. A detailed Executive Summary, covering all the tasks, is also available as Volume 1.

Prediction of Fatigue Crack Growth in Gas Turbine Engine Blades Using Acoustic Emission

PubMed Central

Zhang, Zhiheng; Yang, Guoan; Hu, Kun

2018-01-01

Fatigue failure is the main type of failure that occurs in gas turbine engine blades and an online monitoring method for detecting fatigue cracks in blades is urgently needed. Therefore, in this present study, we propose the use of acoustic emission (AE) monitoring for the online identification of the blade status. Experiments on fatigue crack propagation based on the AE monitoring of gas turbine engine blades and TC11 titanium alloy plates were conducted. The relationship between the cumulative AE hits and the fatigue crack length was established, before a method of using the AE parameters to determine the crack propagation stage was proposed. A method for predicting the degree of crack propagation and residual fatigue life based on the AE energy was obtained. The results provide a new method for the online monitoring of cracks in the gas turbine engine blade. PMID:29693556

Prediction of Fatigue Crack Growth in Gas Turbine Engine Blades Using Acoustic Emission.

PubMed

Zhang, Zhiheng; Yang, Guoan; Hu, Kun

2018-04-25

Fatigue failure is the main type of failure that occurs in gas turbine engine blades and an online monitoring method for detecting fatigue cracks in blades is urgently needed. Therefore, in this present study, we propose the use of acoustic emission (AE) monitoring for the online identification of the blade status. Experiments on fatigue crack propagation based on the AE monitoring of gas turbine engine blades and TC11 titanium alloy plates were conducted. The relationship between the cumulative AE hits and the fatigue crack length was established, before a method of using the AE parameters to determine the crack propagation stage was proposed. A method for predicting the degree of crack propagation and residual fatigue life based on the AE energy was obtained. The results provide a new method for the online monitoring of cracks in the gas turbine engine blade.

Installation of TVC Actuators in a Two Axis Inertial Load Simulator Test Stand

NASA Technical Reports Server (NTRS)

Dziubanek, Adam

2013-01-01

This paper is about the installation of Space Shuttle Main Engines (SSME) actuators in the new Two Axis Inertial Load Simulator (ILS) at MSFC. The new test stand will support the core stage of the Space Launch System (SLS). Because of the unique geometry of the new test stand standard actuator installation procedures will not work. I have been asked to develop a design on how to install the actuators into the new test stand. After speaking with the engineers and technicians I have created a possible design solution. Using Pro Engineer design software and running my own stress calculations I have proven my design is feasible. I have learned how to calculate the stresses my design will see from this task. From the calculations I have learned I have over built the apparatus. I have also expanded my knowledge of Pro Engineer and was able to create a model of my idea.

Saturn Apollo Program

NASA Image and Video Library

1968-01-09

A cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), initially had a thrust of 188,000 pounds each for a combined thrust of over 1,500,000 pounds. Later, the H-1 engine was upgraded to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program. This photo depicts a single modified H-1 engine. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This chart is an illustration of J-2 Engine characteristics. A cluster of five J-2 engines powered the Saturn V S-II (second) stage with each engine providing a thrust of 200,000 pounds. A single J-2 engine powered the S-IVB stage, the Saturn IB second stage, and the Saturn V third stage. The engine was uprated to provide 230,000 pounds of thrust for the fourth Apollo Saturn V flight and subsequent missions. Burning liquid hydrogen as fuel and using liquid oxygen as the oxidizer, the cluster of five J-2 engines for the S-II stage burned over one ton of propellant per second, during about 6 1/2 minutes of operation, to take the vehicle to an altitude of about 108 miles and a speed of near orbital velocity, about 17,400 miles per hour.

SSME structural dynamic model development

NASA Technical Reports Server (NTRS)

Foley, Michael J.

1989-01-01

The high pressure fuel turbopump (HPFTP) is a major component of the Space Shuttle Main Engine (SSME) powerhead. The device is a three stage centrifugal pump that is directly driven by a two stage hot gas turbine. The purpose of the pump is to deliver fuel (liquid hydrogen) from the low pressure fuel turbopump (LPFTP) through the main fuel valve (MFV) to the thrust chamber coolant circuits. In doing so, the pump pressurizes the fuel from an inlet pressure of approximately 178 psi to a discharge pressure of over 6000 psi. At full power level (FPL), the pump rotates at a speed of over 37,000 rpm while generating approximately 77,000 horsepower. Obviously, a pump failure at these speeds and power levels could jeopardize the mission. Results are summarized for work in which the solutions obtained from analytical models of the fuel turbopump impellers are compared with the results obtained from dynamic tests.

High-Payoff Space Transportation Design Approach with a Technology Integration Strategy

NASA Technical Reports Server (NTRS)

McCleskey, C. M.; Rhodes, R. E.; Chen, T.; Robinson, J.

2011-01-01

A general architectural design sequence is described to create a highly efficient, operable, and supportable design that achieves an affordable, repeatable, and sustainable transportation function. The paper covers the following aspects of this approach in more detail: (1) vehicle architectural concept considerations (including important strategies for greater reusability); (2) vehicle element propulsion system packaging considerations; (3) vehicle element functional definition; (4) external ground servicing and access considerations; and, (5) simplified guidance, navigation, flight control and avionics communications considerations. Additionally, a technology integration strategy is forwarded that includes: (a) ground and flight test prior to production commitments; (b) parallel stage propellant storage, such as concentric-nested tanks; (c) high thrust, LOX-rich, LOX-cooled first stage earth-to-orbit main engine; (d) non-toxic, day-of-launch-loaded propellants for upper stages and in-space propulsion; (e) electric propulsion and aero stage control.

Evaluation of SSME test data reduction methods

NASA Technical Reports Server (NTRS)

Santi, L. Michael

1994-01-01

Accurate prediction of hardware and flow characteristics within the Space Shuttle Main Engine (SSME) during transient and main-stage operation requires a significant integration of ground test data, flight experience, and computational models. The process of integrating SSME test measurements with physical model predictions is commonly referred to as data reduction. Uncertainties within both test measurements and simplified models of the SSME flow environment compound the data integration problem. The first objective of this effort was to establish an acceptability criterion for data reduction solutions. The second objective of this effort was to investigate the data reduction potential of the ROCETS (Rocket Engine Transient Simulation) simulation platform. A simplified ROCETS model of the SSME was obtained from the MSFC Performance Analysis Branch . This model was examined and tested for physical consistency. Two modules were constructed and added to the ROCETS library to independently check the mass and energy balances of selected engine subsystems including the low pressure fuel turbopump, the high pressure fuel turbopump, the low pressure oxidizer turbopump, the high pressure oxidizer turbopump, the fuel preburner, the oxidizer preburner, the main combustion chamber coolant circuit, and the nozzle coolant circuit. A sensitivity study was then conducted to determine the individual influences of forty-two hardware characteristics on fourteen high pressure region prediction variables as returned by the SSME ROCETS model.

Technology of interdisciplinary open-ended designing in engineering education

NASA Astrophysics Data System (ADS)

Isaev, A. P.; Plotnikov, L. V.; Fomin, N. I.

2017-11-01

Author’s technology of interdisciplinary open-ended engineering is presented in this article. This technology is an integrated teaching method that significantly increases the practical component in the educational program. Author’s technology creates the conditions to overcome the shortcomings in the engineering education. The basic ideas of the technology of open-ended engineering, experience of their implementation in higher education and the author’s vision of the teaching technology are examined in the article. The main stages of development process of the author’s technology of open-ended engineering to prepare students (bachelor) of technical profile are presented in the article. Complex of the methodological tools and procedures is shown in the article. This complex is the basis of the developed training technology that is used in educational process in higher school of engineering (UrFU). The organizational model of the technology of open-ended engineering is presented. Organizational model integrates the functions in the creation and implementation of all educational program. Analysis of the characteristics of educational activity of students working on author’s technology of interdisciplinary open-ended engineering is presented. Intermediate results of the application of author’s technology in the educational process of the engineering undergraduate are shown.

Study of solid rocket motor for space shuttle booster, volume 2, book 1

NASA Technical Reports Server (NTRS)

1972-01-01

The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

Saturn Apollo Program

NASA Image and Video Library

1963-01-01

This drawing clearly shows the comparative sizes of the rocket engines used to launch the Saturn vehicles. The RL-10 and the H-1 engines were used to launch the Saturn I rockets. The J-2 engine was used on the second stage of Saturn IB and the second and third stages of Saturn V. The F-1 engine was used on the first stage of the Saturn V.

HSCT Sector Combustor Evaluations for Demonstration Engine

NASA Technical Reports Server (NTRS)

Greenfield, Stuart; Heberling, Paul; Kastl, John; Matulaitis, John; Huff, Cynthia

2004-01-01

In LET Task 10, critical development issues of the HSCT lean-burn low emissions combustor were addressed with a range of engineering tools. Laser diagnostics and CFD analysis were applied to develop a clearer understanding of the fuel-air premixing process and premixed combustion. Subcomponent tests evaluated the emissions and operability performance of the fuel-air premixers. Sector combustor tests evaluated the performance of the integrated combustor system. A 3-cup sector was designed and procured for laser diagnostics studies at NASA Glenn. The results of these efforts supported the earlier selection of the Cyclone Swirler as the pilot stage premixer and the IMFH (Integrated Mixer Flame Holder) tube as the main stage premixer of the LPP combustor. In the combustor system preliminary design subtask, initial efforts to transform the sector combustor design into a practical subscale engine combustor met with significant challenges. Concerns about the durability of a stepped combustor dome and the need for a removable fuel injection system resulted in the invention and refinement of the MRA (Multistage Radial Axial) combustor system in 1994. The MRA combustor was selected for the HSR Phase II LPP subscale combustor testing in the CPC Program.

J-2 Engine ready to go into test stand

NASA Technical Reports Server (NTRS)

1965-01-01

Two technicians watch carefully as cables prepare to lift a J-2 engine into a test stand. The J-2 powered the second stage and the third stage of the Saturn V moon rocket. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Second Stage (S-II) Plays Key Role in Apollo missions

NASA Technical Reports Server (NTRS)

1970-01-01

This photograph of the Saturn V Second Stage (S-II) clearly shows the cluster of five powerful J-2 engines needed to boost the Apollo spacecraft into earth orbit following first stage separation. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Computational Fluid Dynamics (CFD) Analysis for the Reduction of Impeller Discharge Flow Distortion

NASA Technical Reports Server (NTRS)

Garcia, R.; McConnaughey, P. K.; Eastland, A.

1993-01-01

The use of Computational Fluid Dynamics (CFD) in the design and analysis of high performance rocket engine pumps has increased in recent years. This increase has been aided by the activities of the Marshall Space Flight Center (MSFC) Pump Stage Technology Team (PSTT). The team's goals include assessing the accuracy and efficiency of several methodologies and then applying the appropriate methodology(s) to understand and improve the flow inside a pump. The PSTT's objectives, team membership, and past activities are discussed in Garcia1 and Garcia2. The PSTT is one of three teams that form the NASA/MSFC CFD Consortium for Applications in Propulsion Technology (McConnaughey3). The PSTT first applied CFD in the design of the baseline consortium impeller. This impeller was designed for the Space Transportation Main Engine's (STME) fuel turbopump. The STME fuel pump was designed with three impeller stages because a two-stage design was deemed to pose a high developmental risk. The PSTT used CFD to design an impeller whose performance allowed for a two-stage STME fuel pump design. The availability of this design would have lead to a reduction in parts, weight, and cost had the STME reached production. One sample of the baseline consortium impeller was manufactured and tested in a water rig. The test data showed that the impeller performance was as predicted and that a two-stage design for the STME fuel pump was possible with minimal risk. The test data also verified another CFD predicted characteristic of the design that was not desirable. The classical 'jet-wake' pattern at the impeller discharge was strengthened by two aspects of the design: by the high head coefficient necessary for the required pressure rise and by the relatively few impeller exit blades, 12, necessary to reduce manufacturing cost. This 'jet-wake pattern produces an unsteady loading on the diffuser vanes and has, in past rocket engine programs, lead to diffuser structural failure. In industrial applications, this problem is typically avoided by increasing the space between the impeller and the diffuser to allow the dissipation of this pattern and, hence, the reduction of diffuser vane unsteady loading. This approach leads to small performance losses and, more importantly in rocket engine applications, to significant increases in the pump's size and weight. This latter consideration typically makes this approach unacceptable in high performance rocket engines.

System Study for Axial Vane Engine Technology

NASA Technical Reports Server (NTRS)

Badley, Patrick R.; Smith, Michael R.; Gould, Cedric O.

2008-01-01

The purpose of this engine feasibility study was to determine the benefits that can be achieved by incorporating positive displacement axial vane compression and expansion stages into high bypass turbofan engines. These positive-displacement stages would replace some or all of the conventional compressor and turbine stages in the turbine engine, but not the fan. The study considered combustion occurring internal to an axial vane component (i.e., Diesel engine replacing the standard turbine engine combustor, burner, and turbine); and external continuous flow combustion with an axial vane compressor and an axial vane turbine replacing conventional compressor and turbine systems.

A Comparison of Propulsion Concepts for SSTO Reusable Launchers

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

This paper discusses the relevant selection criteria for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed for this role against these criteria. The engine types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Liquid Air Cycle Engines. In the authors opinion none of the above engines are able to meet all the necessary criteria for an SSTO propulsion system simultaneously. However by selecting appropriate features from each it is possible to synthesise a new class of engines which are specifically optimised for the SSTO role. The resulting engines employ precooling of the airstream and a high internal pressure ratio to enable a relatively conventional high pressure rocket combustion chamber to be utilised in both airbreathing and rocket modes. This results in a significant mass saving with installation advantages which by careful design of the cycle thermodynamics enables the full potential of airbreathing to be realised. The SABRE engine which powers the SKYLON launch vehicle is an example of one of these so called `Precooled hybrid airbreathing rocket engines' and the concep- tual reasoning which leads to its main design parameters are described in the paper.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

RL-10 engine characteristics. The RL-10 engine was developed under the management of the Marshall Space Flight Center (MSFC) to power the Saturn I upper stage (S-IV stage). The six RL-10 engines, which used liquid hydrogen and liquid oxygen as propellants, were arranged in a circle on the aft end of the S-IV stage.

Red River of the North Main Stem: Technical Information Report (with Special Emphasis on Agricultural Levees).

DTIC Science & Technology

1982-07-01

farmers to construct their own levees. These levees prevented flooding of thousands of acres of cropland. This success led to extensive agricultural...levee construction by farmers on both sides of the river. Various Federal and State agencies expressed concern over the potential adverse impacts of...uncontrolled levee construction . The Corps of Engineer’s analysis showed that continued levee construction would significantly increase flood stages and

The Second Stage of a Saturn V Ready For Test

NASA Technical Reports Server (NTRS)

1970-01-01

This Saturn V S-II (second) stage is being lifted into position for a test at the Vehicle Assembly Building at the Kennedy Space Center. When the Saturn V booster stage (S-IC) burned out and dropped away, power for the Saturn was provided by the 82-foot-long and 33-foot-diameter S-II stage. Developed by the Space Division of North American Aviation under the direction of the Marshall Space Flight Center, the stage utilized five J-2 engines, each producing 200,000 pounds of thrust. The engines used liquid oxygen and liquid hydrogen as propellants. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Noise and Fuel Burn Reduction Potential of an Innovative Subsonic Transport Configuration

NASA Technical Reports Server (NTRS)

Guo, Yueping; Nickol, Craig L.; Thomas, Russell H.

2014-01-01

A study is presented for the noise and fuel burn reduction potential of an innovative double deck concept aircraft with two three-shaft direct-drive turbofan engines. The engines are mounted from the fuselage so that the engine inlet is over the main wing. It is shown that such an aircraft can achieve a cumulative Effective Perceived Noise Level (EPNL) about 28 dB below the current aircraft noise regulations of Stage 4. The combination of high bypass ratio engines and advanced wing design with laminar flow control technologies provide fuel burn reduction and low noise levels simultaneously. For example, the fuselage mounted engine position provides more than 4 EPNLdB of noise reduction by shielding the inlet radiated noise. To identify the potential effect of noise reduction technologies on this concept, parametric studies are presented to reveal the system level benefits of various emerging noise reduction concepts, for both engine and airframe noise reduction. These concepts are discussed both individually to show their respective incremental noise reduction potential and collectively to assess their aggregate effects on the total noise. Through these concepts approximately about 8 dB of additional noise reduction is possible, bringing the cumulative noise level of this aircraft to 36 EPNLdB below Stage 4, if the entire suite of noise reduction technologies would mature to practical application. In a final step, an estimate is made for this same aircraft concept but with higher bypass ratio, geared, turbofan engines. With this geared turbofan propulsion system, the noise is estimated to reach as low as 40-42 dB below Stage 4 with a fuel burn reduction of 43-47% below the 2005 best-in-class aircraft baseline. While just short of the NASA N+2 goals of 42 dB and 50% fuel burn reduction, for a 2025 in service timeframe, this assessment shows that this innovative concept warrants refined study. Furthermore, this design appears to be a viable potential future passenger aircraft, not only in meeting the regulatory requirements, but also in competing with aircraft of different advanced designs within this N+2 timeframe and goal framework.

Optimization of Turbine Rim Seals

NASA Technical Reports Server (NTRS)

Wagner, J. H.; Tew, D. E.; Stetson, G. M.; Sabnis, J. S.

2006-01-01

Experiments are being conducted to gain an understanding of the physics of rim scale cavity ingestion in a turbine stage with the high-work, single-stage characteristics envisioned for Advanced Subsonic Transport (AST) aircraft gas turbine engines fo the early 21st century. Initial experimental measurements to be presented include time-averaged turbine rim cavity and main gas path static pressure measurements for rim seal coolant to main gas path mass flow ratios between 0 and 0.02. The ultimate objective of this work is develop improved rim seal design concepts for use in modern high-work, single sage turbines n order to minimize the use of secondary coolant flow. Toward this objective the time averaged and unsteady data to be obtained in these experiments will be used to 1) Quantify the impact of the rim cavity cooling air on the ingestion process. 2) Quantify the film cooling benefits of the rim cavity purge flow in the main gas path. 3) Quantify the impact of the cooling air on turbine efficiency. 4) Develop/evaluate both 3D CFD and analytical models of the ingestion/cooling process.

[Development of Web-based multimedia content for a physical examination and health assessment course].

PubMed

Oh, Pok-Ja; Kim, Il-Ok; Shin, Sung-Rae; Jung, Hoe-Kyung

2004-10-01

This study was to develop Web-based multimedia content for Physical Examination and Health Assessment. The multimedia content was developed based on Jung's teaching and learning structure plan model, using the following 5 processes : 1) Analysis Stage, 2) Planning Stage, 3) Storyboard Framing and Production Stage, 4) Program Operation Stage, and 5) Final Evaluation Stage. The web based multimedia content consisted of an intro movie, main page and sub pages. On the main page, there were 6 menu bars that consisted of Announcement center, Information of professors, Lecture guide, Cyber lecture, Q&A, and Data centers, and a site map which introduced 15 week lectures. In the operation of web based multimedia content, HTML, JavaScript, Flash, and multimedia technology (Audio and Video) were utilized and the content consisted of text content, interactive content, animation, and audio & video. Consultation with the experts in context, computer engineering, and educational technology was utilized in the development of these processes. Web-based multimedia content is expected to offer individualized and tailored learning opportunities to maximize and facilitate the effectiveness of the teaching and learning process. Therefore, multimedia content should be utilized concurrently with the lecture in the Physical Examination and Health Assessment classes as a vital teaching aid to make up for the weakness of the face-to- face teaching-learning method.

Configuration evaluation and criteria plan. Volume 2: Evaluation critera plan (preliminary). Space Transportation Main Engine (STME) configuration study

NASA Technical Reports Server (NTRS)

Bair, E. K.

1986-01-01

The unbiased selection of the Space Transportation Main Engine (STME) configuration requires that the candidate engines be evaluated against a predetermined set of criteria which must be properly weighted to emphasize critical requirements defined prior to the actual evaluation. The evaluation and selection process involves the following functions: (1) determining if a configuration can satisfy basic STME requirements (yes/no); (2) defining the evaluation criteria; (3) selecting the criteria relative importance or weighting; (4) determining the weighting sensitivities; and (5) establishing a baseline for engine evaluation. The criteria weighting and sensitivities are cost related and are based on mission models and vehicle requirements. The evaluation process is used as a coarse screen to determine the candidate engines for the parametric studies and as a fine screen to determine concept(s) for conceptual design. The criteria used for the coarse and fine screen evaluation process is shown. The coarse screen process involves verifying that the candidate engines can meet the yes/no screening requirements and a semi-subjective quantitative evaluation. The fine screen engines have to meet all of the yes/no screening gates and are then subjected to a detailed evaluation or assessment using the quantitative cost evaluation processes. The option exists for re-cycling a concept through the quantitative portion of the screening and allows for some degree of optimization. The basic vehicle is a two stage LOX/HC, LOX/LH2 parallel burn vehicle capable of placing 150,000 lbs in low Earth orbit (LEO).

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Astrophysics Data System (ADS)

Manski, Detlef; Martin, James A.

1988-07-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Manski, Detlef; Martin, James A.

1988-01-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Effects of nozzle-strut integrated design concepton on the subsonic turbine stage flowfield

NASA Astrophysics Data System (ADS)

Liu, Jun; Du, Qiang; Liu, Guang; Wang, Pei; Zhu, Junqiang

2014-10-01

In order to shorten aero-engine axial length, substituting the traditional long chord thick strut design accompanied with the traditional low pressure(LP) stage nozzle, LP turbine is integrated with intermediate turbine duct (ITD). In the current paper, five vanes of the first stage LP turbine nozzle is replaced with loaded struts for supporting the engine shaft, and providing oil pipes circumferentially which fulfilled the areo-engine structure requirement. However, their bulky geometric size represents a more effective obstacle to flow from high pressure (HP) turbine rotor. These five struts give obvious influence for not only the LP turbine nozzle but also the flowfield within the ITD, and hence cause higher loss. Numerical investigation has been undertaken to observe the influence of the Nozzle-Strut integrated design concept on the flowfield within the ITD and the nearby nozzle blades. According to the computational results, three main conclusions are finally obtained. Firstly, a noticeable low speed area is formed near the strut's leading edge, which is no doubt caused by the potential flow effects. Secondly, more severe radial migration of boundary layer flow adjacent to the strut's pressure side have been found near the nozzle's trailing edge. Such boundary layer migration is obvious, especially close to the shroud domain. Meanwhile, radial pressure gradient aggravates this phenomenon. Thirdly, velocity distribution along the strut's pressure side on nozzle's suction surface differs, which means loading variation of the nozzle. And it will no doubt cause nonuniform flowfield faced by the downstream rotor blade.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

Workmen inspect a J-2 engine at Rocketdyne's Canoga Park, California production facility. The J-2, developed under the direction of the Marshall Space Flight Center, was propelled by liquid hydrogen and liquid oxygen. A single J-2 engine was used in the S-IVB stage (the second stage of the Saturn IB and third stage for the Saturn V) and a cluster of five J-2 engines was used to propel the second stage of the Saturn V, the S-II. Initially rated at 200,000 pounds of thrust, the J-2 engine was later uprated in the Saturn V program to 230,000 pounds.

Design Evolution and Verification of the A-3 Chemical Steam Generator

NASA Technical Reports Server (NTRS)

Kirchner, Casey K.

2009-01-01

Following is an overview of the Chemical Steam Generator system selected to provide vacuum conditions for a new altitude test facility, the A-3 Test Stand at Stennis Space Center (SSC) in Bay St. Louis, MS. A-3 will serve as NASA s primary facility for altitude testing of the J-2X rocket engine, to be used as the primary propulsion device for the upper stages of the Ares launch vehicles. The Chemical Steam Generators (CSGs) will produce vacuum conditions in the test cell through the production and subsequent supersonic ejection of steam into a diffuser downstream of the J-2X engine nozzle exit. The Chemical Steam Generators chosen have a rich heritage of operation at rocket engine altitude test facilities since the days of the Apollo program and are still in use at NASA White Sands Test Facility (WSTF) in New Mexico. The generators at WSTF have been modified to a degree, but are still very close to the heritage design. The intent for the A-3 implementation is to maintain this heritage design as much as possible, making minimal updates only where necessary to substitute for obsolete parts and to increase reliability. Reliability improvements are especially desired because the proposed system will require 27 generators, which is nine times the largest system installed in the 1960s. Improvements were suggested by the original design firm, Reaction Motors, by NASA SSC and NASA WSTF engineers, and by the A-3 test stand design contractor, Jacobs Technology, Inc. (JTI). This paper describes the range of improvements made to the design to date, starting with the heritage generator and the minor modifications made over time at WSTF, to the modernized configuration which will be used at A-3. The paper will discuss NASA s investment in modifications to SSC s E-2 test facility fire a full-scale Chemical Steam Generator in advance of the larger steam system installation at A-3. Risk mitigation testing will be performed in early 2009 at this test facility to verify that the CSGs operate as expected. The generator which will undergo this testing is of the most recent A-3 configuration, and will be instrumented far in excess of what is normally required for operation. The extra data will allow for easier troubleshooting and more complete knowledge of expected generator performance. In addition, the early testing will give SSC personnel experience in operating the CSG systems, which will expedite the process of installation and activation at A-3. Each Chemical Steam Generator is supported by a complement of valves, instruments, and flow control devices, with the entire assembly called a "module." The generators will be installed in groups of three, historically called "units". A module is so called because of its modular ability to be replaced or serviced without disturbing the other two modules installed on the same unit. A module is pictured in Figure 1, shown with its generator secured by white bands in its shipping (vs. installed) configuration. The heritage system at WSTF is composed of a single unit (three generator modules), pictured in Figure 2 as it was installed in 1965. In contrast, A-3 will have nine units operating in parallel to achieve vacuum conditions appropriate for testing the J-2X engine. Each of the combustors operates in two modes and achieves the so-called "full-steam" mode after all three of its stages ignite. Ignition of the first stage is achieved by exciting a spark plug; the second stage and main stage are lit by the flame front of the previous stage. The main stage burns approximately 97% of the total propellant flow and uses the heat energy to vaporize water into superheated steam. While the main stage remains unlit, the combustor is in so-called "idle" mode. In the WSTF system, this idle mode is not optimized for water usage, and does not need to be, as the water is pumped from a large reservoir. The water supply at A-3 will be contained in tanks with finite volume, so water optimization is preferred for the modnized configuration. Multiple solutions for this issue have been proposed, with the leading concept being a change to the operational definition of "idle mode," with the generator running in a lower heat flux condition.

One output function: a misconception of students studying digital systems - a case study

NASA Astrophysics Data System (ADS)

Trotskovsky, E.; Sabag, N.

2015-05-01

Background:Learning processes are usually characterized by students' misunderstandings and misconceptions. Engineering educators intend to help their students overcome their misconceptions and achieve correct understanding of the concept. This paper describes a misconception in digital systems held by many students who believe that combinational logic circuits should have only one output. Purpose:The current study aims to investigate the roots of the misconception about one-output function and the pedagogical methods that can help students overcome the misconception. Sample:Three hundred and eighty-one students in the Departments of Electrical and Electronics and Mechanical Engineering at an academic engineering college, who learned the same topics of a digital combinational system, participated in the research. Design and method:In the initial research stage, students were taught according to traditional method - first to design a one-output combinational logic system, and then to implement a system with a number of output functions. In the main stage, an experimental group was taught using a new method whereby they were shown how to implement a system with several output functions, prior to learning about one-output systems. A control group was taught using the traditional method. In the replication stage (the third stage), an experimental group was taught using the new method. A mixed research methodology was used to examine the results of the new learning method. Results:Quantitative research showed that the new teaching approach resulted in a statistically significant decrease in student errors, and qualitative research revealed students' erroneous thinking patterns. Conclusions:It can be assumed that the traditional teaching method generates an incorrect mental model of the one-output function among students. The new pedagogical approach prevented the creation of an erroneous mental model and helped students develop the correct conceptual understanding.

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.

Heat transfer and pressure measurements and comparison with prediction for the SSME two-stage turbine

NASA Technical Reports Server (NTRS)

Dunn, M. G.; Kim, J.

1992-01-01

Time averaged Stanton number and surface pressure distributions are reported for the first stage vane row, the first stage blade row, and the second stage vane row of the Rocketdyne Space Shuttle Main Engine (SSME) two-stage fuel-side turbine. Unsteady pressure envelope measurements for the first blade are also reported. These measurements were made at 10 percent, 50 percent, and 90 percent span on both the pressure and suction surfaces of the first stage components. Additional Stanton number measurements were made on the first stage blade platform, blade tip, and shroud, and at 50 percent span on the second vane. A shock tube was used as a short duration source of heated and pressurized air to which the turbine was subjected. Platinum thin film heat flux gages were used to obtain the heat flux measurements, while miniature silicon diaphragm flush-mounted pressure transducers were used to obtain the pressure measurements. The first stage vane Stanton number distributions are compared with predictions obtained using a version of STAN5 and quasi-3D Navier-Stokes solution. This same quasi-3D N-S code was also used to obtain predictions for the first blade and the second vane.

Mathematical modeling and optimization of flow structure in stage of francis turbine of micro gas turbine power plant

NASA Astrophysics Data System (ADS)

Kartashev, A. L.; Vaulin, S. D.; Kartasheva, M. A.; Martynov, A. A.; Safonov, E. V.

2016-06-01

This article presents information about the main distinguishing features of microturbine power plants. The justification of the use of Francis turbine in microturbine power plants with rated power of 100 kW is given. Initial analytical engineering calculations of the turbine (without using computational fluid dynamics) with appropriate calculation methods are considered. The parametric study of nozzle blade and whole turbine stage using ANSYS CFX is descripted. The calculations determined the optimal geometry on the criterion of maximizing efficiency at total pressure ratio. The calculation results are presented in graphical form, as well as the velocity and pressure fields at the interscapular channels of nozzle unit and the impeller.

Blown Away

NASA Astrophysics Data System (ADS)

Scott, Alan

2014-04-01

On Aug. 13, 2011, at 8:45 p.m. country music fans were eagerly awaiting the band Sugarland to make its entry onto the main stage at the Indiana State Fairgrounds. Also competing for the fans' attention was an approaching storm. Sugarland never made their entrance. At 8:49 p.m. the stage rigging was hit by 59 mile/h (94 km/h) winds causing it to collapse, killing seven people and injuring 58. A simplified facsimile model of this collapse is presented here that can be used as a case study in applying fundamental physics concepts appropriate for an introductory class. It is a forensic physics/engineering investigation. Afterwards, key aspects of the actual investigation will be presented.

NASA's Space Launch System: Development and Progress

NASA Technical Reports Server (NTRS)

Honeycutt, John; Lyles, Garry

2016-01-01

NASA is embarked on a new era of space exploration that will lead to new capabilities, new destinations, and new discoveries by both human and robotic explorers. Today, the International Space Station (ISS), supported by NASA's commercial partners, and robotic probes, are yielding knowledge that will help make this exploration possible. NASA is developing both the Orion crew vehicle and the Space Launch System (SLS) that will carry out a series of increasingly challenging missions that will eventually lead to human exploration of Mars. This paper will discuss the development and progress on the SLS. The SLS architecture was designed to be safe, affordable, and sustainable. The current configuration is the result of literally thousands of trade studies involving cost, performance, mission requirements, and other metrics. The initial configuration of SLS, designated Block 1, will launch a minimum of 70 metric tons (t) into low Earth orbit - significantly greater capability than any current launch vehicle. It is designed to evolve to a capability of 130 t through the use of upgraded main engines, advanced boosters, and a new upper stage. With more payload mass and volume capability than any rocket in history, SLS offers mission planners larger payloads, faster trip times, simpler design, shorter design cycles, and greater opportunity for mission success. Since the program was officially created in fall 2011, it has made significant progress toward first launch readiness of the Block 1 vehicle in 2018. Every major element of SLS continued to make significant progress in 2015. The Boosters element fired Qualification Motor 1 (QM-1) in March 2015, to test the 5-segment motor, including new insulation, joint, and propellant grain designs. The Stages element marked the completion of more than 70 major components of test article and flight core stage tanks. The Liquid Engines element conducted seven test firings of an RS-25 engine under SLS conditions. The Spacecraft/Payload Integration and Evolution element marked completion of the upper stage test article. Major work continues in 2016 as the program continues both flight and development RS-25 engine testing, begins welding test article and flight core stage tanks, completes stage adapter manufacturing, and test fires the second booster qualification motor. This paper will discuss the program's key accomplishments to date and the challenging work ahead for what will be the world's most capable launch vehicle.

Second Stage (S-II) Arrives at Marshall Space Flight Center For Testing

NASA Technical Reports Server (NTRS)

2004-01-01

The business end of a Second Stage (S-II) slowly emerges from the shipping container as workers prepare to transport the Saturn V component to the testing facility at MSFC. The Second Stage (S-II) underwent vibration and engine firing tests. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1970-01-22

This Saturn V S-II (second) stage is being lifted into position for a test at the Vehicle Assembly Building at the Kennedy Space Center. When the Saturn V booster stage (S-IC) burned out and dropped away, power for the Saturn was provided by the 82-foot-long and 33-foot-diameter S-II stage. Developed by the Space Division of North American Aviation under the direction of the Marshall Space Flight Center, the stage utilized five J-2 engines, each producing 200,000 pounds of thrust. The engines used liquid oxygen and liquid hydrogen as propellants. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

40 CFR 92.215 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2010 CFR

2010-07-01

... stage to the production stage, or to locomotive (or engine) manufacture, remanufacture, or assembly is...) In the case where a current production engine is modified for use as a certification engine or in a... not derived from a current production engine, a general description of the buildup of the engine (e.g...

From Paper to Production: An Update on NASA's Upper Stage Engine for Exploration

NASA Technical Reports Server (NTRS)

Kynard, Mike

2010-01-01

In 2006, NASA selected an evolved variant of the proven Saturn/Apollo J-2 upper stage engine to power the Ares I crew launch vehicle upper stage and the Ares V cargo launch vehicle Earth departure stage (EDS) for the Constellation Program. Any design changes needed by the new engine would be based where possible on proven hardware from the Space Shuttle, commercial launchers, and other programs. In addition to the thrust and efficiency requirements needed for the Constellation reference missions, it would be an order of magnitude safer than past engines. It required the J-2X government/industry team to develop the highest performance engine of its type in history and develop it for use in two vehicles for two different missions. In the attempt to achieve these goals in the past five years, the Upper Stage Engine team has made significant progress, successfully passing System Requirements Review (SRR), System Design Review (SDR), Preliminary Design Review (PDR), and Critical Design Review (CDR). As of spring 2010, more than 100,000 experimental and development engine parts have been completed or are in various stages of manufacture. Approximately 1,300 of more than 1,600 engine drawings have been released for manufacturing. This progress has been due to a combination of factors: the heritage hardware starting point, advanced computer analysis, and early heritage and development component testing to understand performance, validate computer modeling, and inform design trades. This work will increase the odds of success as engine team prepares for powerpack and development engine hot fire testing in calendar 2011. This paper will provide an overview of the engine development program and progress to date.

Lessons Learned from the Design, Certification, and Operation of the Space Shuttle Integrated Main Propulsion System (IMPS)

NASA Technical Reports Server (NTRS)

Martinez, Hugo E.; Albright, John D.; D'Amico, Stephen J.; Brewer, John M.; Melcher, John C., IV

2011-01-01

The Space Shuttle Integrated Main Propulsion System (IMPS) consists of the External Tank (ET), Orbiter Main Propulsion System (MPS), and Space Shuttle Main Engines (SSMEs). The IMPS is tasked with the storage, conditioning, distribution, and combustion of cryogenic liquid hydrogen (LH2) and liquid oxygen (LO2) propellants to provide first and second stage thrust for achieving orbital velocity. The design, certification, and operation of the associated IMPS hardware have produced many lessons learned over the course of the Space Shuttle Program (SSP). A subset of these items will be discussed in this paper for consideration when designing, building, and operating future spacecraft propulsion systems. This paper will focus on lessons learned related to Orbiter MPS and is the first of a planned series to address the subject matter.

[Cell engineering in nephrology: The current state and prospectives from the point of view of military medicine].

PubMed

Bel'skih, A N; Golota, A S; Krassii, A B; Nagibovich, O A

2015-09-01

This article is dedicated to the current state and prospectives of cell engineering in nephrology from the point of view of military medicine. The review is based on publications from January 1, 2014, to June 1, 2015, and consists of two parts. In the first part the main directions of the cell engineering development are mentioned. The only two clinical trials existed in the field are discussed in more detail. The second part deals with prospectives of cell engineering in nephrology. It is shown that currently this field is in the stage of preclinical experimentation. Of the two known clinical trials the first has failed to demonstrate any effectiveness of cell engineering, the second--will be completed only at the end of 2016. Also, the review notes an extraordinary cost of cell engineering experiments in nephrology. The analysis of publications allows to come to a conclusion that the future progress in prevention and treatment of acute kidney injury could go not in the direction of cell engineering but rather non-cell technologies. One of. the practical consequence of such a conclusion is a necessity to continue the improvement of already existing methods of machine renal replacement therapy.

Sensitivity of outflow peaks and flood stages to the selection of dam breach parameters and simulation models

NASA Astrophysics Data System (ADS)

Singh, Krishan P.; Snorrason, Arni

1984-02-01

Important breach parameters were identified and their ranges were estimated from a detailed study of historical earthdam failures due to overtopping. The U.S. Army Corps of Engineers Hydrologic Engineering Center (HEC) and the National Weather Service (NWS) dam breach models were chosen for evaluation and simulation. Both models use similar input data and breach descriptions, but the HEC uses the hydrologic routing method (modified Puls method), whereas the NWS uses the St. Vénant equations for routing. Information on eight dams in Illinois was taken from the Corps of Engineers inspection reports, and surveyed cross-sections of the downstream channels were supplied by the Division of Water Resources of the Illinois Department of Transportation. Various combinations of breach parameters (failure time, TF; depth of overtopping, hf; and breach size, B) were used for breach simulations by both methods with the 1.00PMF, 0.50PMF and 0.25PMF (probable maximum flood) inflow hydrographs. In general, the flood stage profiles predicted by the NWS were smoother and more reasonable than those predicted by the HEC. For channels with relatively steep slopes, the methods compared fairly well, whereas for the channels with mild slope, the HEC model often predicted oscillating, erratic flood stages, mainly due to its inability to route flood waves satisfactorily in non-prismatic channels. The breach outflow peaks are affected significantly by B but less so by hf. The ratio of outflow peak to inflow peak and the effect of TF on outflow decrease as the drainage area above the dam and impounded storage increase. Flood stage profiles predicted with cross-sections taken from 7.5' maps compared favorably with those predicted using surveyed cross-sections. For the range of breach parameters studied, the range of outflow peaks and flood stages downstream from the dam can be determined for regulatory and disaster prevention measures.

Saturn Apollo Program

NASA Image and Video Library

1965-04-26

Two technicians watch carefully as cables prepare to lift a J-2 engine into a test stand. The J-2 powered the second stage and the third stage of the Saturn V moon rocket. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Real application of BIM in the engineering system design for energy management

NASA Astrophysics Data System (ADS)

Pelipenko, Alexey; Gogina, Elena

2017-10-01

In the article, the information modelling technology (BIM) that is gaining popularity in Russia and in the world, is considered. Its growing relevance relates to many factors: first, attention to this technology by the local and federal authorities; Secondly, with the desire to improve the quality of design documentation, to obtain the correct volumes of materials and equipment; Thirdly, with the tendency to create “smart” cities and, as a result, the rational use of energy resources. Within the framework of this article, on an example of an urban infrastructure object, the pros and cons of this technology were considered. As a facility, a local wastewater treatment plant was chosen. The stages of creating an information model on the available documentation are described: 4 main milestones that need to be implemented. In addition, further possible ways of using the model are described. Presented are the pros and cons of using this technology. Among the main advantages is the possibility of using this information model in the operation of treatment plants and further obtaining actual data for monitoring the condition of equipment and, therefore, controlling the consumable resources; At an early stage, a reduction in the number of mutual intersections of engineering systems; Obtaining the correct specifications. The results of the work described in the article can be used in the following areas: utilities, energy management, design and construction.

The Development of Titanium Alloys for Application in the Space Shuttle Main Engine

NASA Technical Reports Server (NTRS)

Halchak, John A.; Jerman, Gregory A.; Zimmerman, Frank R.

2010-01-01

The high-strength-to-weight ratio of titanium alloys, particularly at cryogenic temperatures, make them attractive for application in rocket engines - offering the potential of superior performance while minimizing component weight. This was particularly attractive for rotating components, such as pump impellers, where titanium alloys presented the potential to achieve a major advance in rotational tip speed, with a reduction in stages and resultant saving in pump weight and complexity. The investigation into titanium alloys for application in cryogenic turbopumps began in the early 1960's. However, it was found that the reactivity of titanium limited applications and produced unique processing challenges. Specialized chemical compositions and processing techniques had to be developed. A substantial amount of material properties testing and trials in experimental turbopumps occurred, ultimately leading to application in the Space Shuttle Main Engine. One particular alloy stood out for use at liquid hydrogen temperatures, Ti-5Al-2.5Sn ELI. This alloy was employed for several critical components. This presentation deals with the development effort, the challenges that were encountered and operational experiences with Ti-5Al-2.5Sn ELI in the SSME.

Technician Works on a Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1977-02-21

A technician prepares a 2.25 percent scale model of the space shuttle for a base heat study in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. This space shuttle project, begun here in July 1976, was aimed at evaluating base heating and pressure prior to the Shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Engineers needed to understand these issues in order to design proper thermal protection. The test’s specific objectives were to measure the heat transfer and pressure distributions around the orbiter’s external tank and solid rocket afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions caused by rocket exhaust-induced separation, and determine gas recovery temperatures using gas temperature probes and heated base components. The shuttle model’s main engines and solid rockets were first fired and then just the main engines to simulate a launch during the testing. Lewis researchers conducted 163 runs in the 10- by 10 during the test program.

Space shuttle main engine definition (phase B). Volume 5: Valves and interconnects. [for space shuttle

NASA Technical Reports Server (NTRS)

Schultz, D. F.

1971-01-01

The steady state thermodynamic cycle balance of the single preburner staged combustion engine, coupled with dynamic transient analyses, dictated in detail the location and requirements for each valve defined in this volume. Valve configuration selections were influenced by overall engine and vehicle system weight and failure mode determinations. Modulating valve actuators are external to the valve and are line replaceable. Development and satisfactory demonstration of a high pressure dynamic shaft seal has made this configuration practical. Pneumatic motor driven actuators that use engine pumped hydrogen gas as the working fluid are used. The helium control system is proposed as a module containing a cluster of solenoid actuated valves. The separable couplings and flanges are designed to assure minimum leakage with minimum coupling weight. The deflection of the seal surface in the flange is defined by finite element analysis that has been confirmed with test data. The seal design proposed has passed preliminary pressure cycling and thermal cycling tests.

Tripropellant engine study

NASA Technical Reports Server (NTRS)

Wheeler, D. B.

1977-01-01

Work conducted was devoted to three main tasks. Thermochemical equilibrium performance data were assembled to establish the expected performance calculations of the mode 1 engine propellant combinations and thermodynamic and transport data for the products of combustion. Turbine drive gas characteristics were also established. Thrust chamber and nozzle cooling studies were devoted to the evaluation of H2, C3H8, CH4, and RP-1 as coolants in the existing SSME cooling circuit geometry. It was found that all these candidate coolants are feasible without limiting the desired operating conditions with the exception of RP-1, which would limit the maximum P(c) to 2000 psia. RP-1 could be used, however, to cool the nozzle only without imposing the chamber pressure limit. A total of 15 candidate engine system cycles were selected and a preliminary engine system balance was conducted for 12 of these systems to establish component operating flowrates, pressures and temperatures. It was found that the staged combustion cycles employing fuel rich LOX/hydrocarbon turbine drive gases are power limited.

Source localization of turboshaft engine broadband noise using a three-sensor coherence method

NASA Astrophysics Data System (ADS)

Blacodon, Daniel; Lewy, Serge

2015-03-01

Turboshaft engines can become the main source of helicopter noise at takeoff. Inlet radiation mainly comes from the compressor tones, but aft radiation is more intricate: turbine tones usually are above the audible frequency range and do not contribute to the weighted sound levels; jet is secondary and radiates low noise levels. A broadband component is the most annoying but its sources are not well known (it is called internal or core noise). Present study was made in the framework of the European project TEENI (Turboshaft Engine Exhaust Noise Identification). Its main objective was to localize the broadband sources in order to better reduce them. Several diagnostic techniques were implemented by the various TEENI partners. As regards ONERA, a first attempt at separating sources was made in the past with Turbomeca using a three-signal coherence method (TSM) to reject background non-acoustic noise. The main difficulty when using TSM is the assessment of the frequency range where the results are valid. This drawback has been circumvented in the TSM implemented in TEENI. Measurements were made on a highly instrumented Ardiden turboshaft engine in the Turbomeca open-air test bench. Two engine powers (approach and takeoff) were selected to apply TSM. Two internal pressure probes were located in various cross-sections, either behind the combustion chamber (CC), the high-pressure turbine (HPT), the free-turbine first stage (TL), or in four nozzle sections. The third transducer was a far-field microphone located around the maximum of radiation, at 120° from the intake centerline. The key result is that coherence increases from CC to HPT and TL, then decreases in the nozzle up to the exit. Pressure fluctuations from HPT and TL are very coherent with the far-field acoustic spectra up to 700 Hz. They are thus the main acoustic source and can be attributed to indirect combustion noise (accuracy decreases above 700 Hz because coherence is lower, but far-field sound spectra also are much lower above 700 Hz).

75 FR 12661 - Airworthiness Directives; General Electric Company CF6-45 and CF6-50 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-17

... borescope inspections of the high-pressure turbine (HPT) rotor stage 1 and stage 2 blades for wear and... already been done. Borescope Inspection of High-Pressure Turbine (HPT) Stage 1 and Stage 2 Rotor Blades (f... turbofan engines with certain low-pressure turbine (LPT) stage 3 disks installed. This AD requires...

Novel 3-D free-form surface profilometry for reverse engineering

NASA Astrophysics Data System (ADS)

Chen, Liang-Chia; Huang, Zhi-Xue

2005-01-01

This article proposes an innovative 3-D surface contouring approach for automatic and accurate free-form surface reconstruction using a sensor integration concept. The study addresses a critical problem in accurate measurement of free-form surfaces by developing an automatic reconstruction approach. Unacceptable measuring accuracy issues are mainly due to the errors arising from the use of inadequate measuring strategies, ending up with inaccurate digitised data and costly post-data processing in Reverse Engineering (RE). This article is thus aimed to develop automatic digitising strategies for ensuring surface reconstruction efficiency, as well as accuracy. The developed approach consists of two main stages, namely the rapid shape identification (RSI) and the automated laser scanning (ALS) for completing 3-D surface profilometry. This developed approach effectively utilises the advantages of on-line geometric information to evaluate the degree of satisfaction of user-defined digitising accuracy under a triangular topological patch. An industrial case study was used to attest the feasibility of the approach.

SSME to RS-25: Challenges of Adapting a Heritage Engine to a New Vehicle Architecture

NASA Technical Reports Server (NTRS)

Ballard, Richard O.

2015-01-01

Following the cancellation of the Constellation program and retirement of the Space Shuttle, NASA initiated the Space Launch System (SLS) program to provide next-generation heavy lift cargo and crew access to space. A key constituent of the SLS architecture is the RS-25 engine, also known as the Space Shuttle Main Engine (SSME). The RS-25 was selected to serve as the main propulsion system for the SLS core stage in conjunction with the solid rocket boosters. This selection was largely based on the maturity and extensive experience gained through 135 missions, 3000+ ground tests, and over a million seconds total accumulated hot-fire time. In addition, there were also over a dozen functional flight assets remaining from the Space Shuttle program that could be leveraged to support the first four flights. However, while the RS-25 is a highly mature system, simply unbolting it from the Space Shuttle boat-tail and installing it on the new SLS vehicle is not a "plug-and-play" operation. In addition to numerous technical integration details involving changes to significant areas such as the environments, interface conditions, technical performance requirements, operational constraints and so on, there were other challenges to be overcome in the area of replacing the obsolete engine control system (ECS). While the magnitude of accomplishing this effort was less than that needed to develop and field a new clean-sheet engine system, the path to the first flight of SLS has not been without unexpected challenges.

Visualization of bacteriophage P1 infection by cryo-electron tomography of tiny Escherichia coli

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

Liu Jun, E-mail: Jun.Liu.1@uth.tmc.edu; Chen Chengyen; Shiomi, Daisuke

2011-09-01

Bacteriophage P1 has a contractile tail that targets the conserved lipopolysaccharide on the outer membrane surface of the host for initial adsorption. The mechanism by which P1 DNA enters the host cell is not well understood, mainly because the transient molecular interactions between bacteriophage and bacteria have been difficult to study by conventional approaches. Here, we engineered tiny E. coli host cells so that the initial stages of P1-host interactions could be captured in unprecedented detail by cryo-electron tomography. Analysis of three-dimensional reconstructions of frozen-hydrated specimens revealed three predominant configurations: an extended tail stage with DNA present in the phagemore » head, a contracted tail stage with DNA, and a contracted tail stage without DNA. Comparative analysis of various conformations indicated that there is uniform penetration of the inner tail tube into the E. coli periplasm and a significant movement of the baseplate away from the outer membrane during tail contraction.« less

Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

NASA Technical Reports Server (NTRS)

Thorpe, Douglas G.

1991-01-01

An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

Technology requirements for advanced earth-orbital transportation systems, dual-mode propulsion

NASA Technical Reports Server (NTRS)

Haefeli, R. C.; Littler, E. G.; Hurley, J. B.; Winter, M. G.

1977-01-01

The application of dual-mode propulsion concepts to fully reusable single-stage-to-orbit (SSTO) vehicles is discussed. Dual-mode propulsion uses main rocket engines that consume hydrocarbon fuels as well as liquid hydrogen fuel. Liquid oxygen is used as the oxidizer. These engine concepts were integrated into transportation vehicle designs capable of vertical takeoff, delivering a payload to earth orbit, and return to earth with a horizontal landing. Benefits of these vehicles were assessed and compared with vehicles using single-mode propulsion (liquid hydrogen and oxygen engines). Technology requirements for such advanced transportation systems were identified. Figures of merit, including life-cycle cost savings and research costs, were derived for dual-mode technology programs, and were used for assessments of potential benefits of proposed technology activities. Dual-mode propulsion concepts display potential for significant cost and performance benefits when applied to SSTO vehicles.

Characterization of ignition transient processes in kerosene-fueled model scramjet engine by dual-pulse laser-induced plasma

NASA Astrophysics Data System (ADS)

Li, Xipeng; Liu, Weidong; Pan, Yu; Yang, Leichao; An, Bin; Zhu, Jiajian

2018-03-01

Dual-pulse laser-induced plasma ignition of kerosene in cavity at model scramjet engine is studied. The simulated flight condition is Ma 6 at 30 km, and the isolator entrance has a Mach number of 2.92, a total pressure of 2.6 MPa and a stagnation temperature of 1650 K. Two independent laser pulses at 532 nm with a pulse width of 10 ns, a diameter of 12 mm and a maximum energy of 300 mJ are focused into cavity for ignition. The flame structure and propagation during transient ignition processes are captured by simultaneous CH* and OH* chemiluminescence imaging. The entire ignition process of kerosene can be divided into five stages, which are referred as turbulent dissipation stage, quasi-stable state, combustion enhancement stage, reverting stage and combustion stabilization stage. A local closed loop of propagations of the burning mixtures from the shear layer into the recirculation zone of cavity is revealed, which the large-scale eddy in the shear layer plays a key role. The enhancement of mass exchange between shear layer and the recirculation zone of cavity could promote the flame propagation process and enhance the ignition capability as well as extend the ignition limits. A cavity shear-layer stabilized combustion of kerosene is established in the supersonic flow roughly 3.3 ms after the laser pulse. Chemical reactions mainly occur in the shear layer and the near-wall zone downstream of the cavity. The distribution of OH* is thicker than CH* at stable combustion condition.

Linear aerospike engine. [for reusable single-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

Kirby, F. M.; Martinez, A.

1977-01-01

A description is presented of a dual-fuel modular split-combustor linear aerospike engine concept. The considered engine represents an approach to an integrated engine for a reusable single-stage-to-orbit (SSTO) vehicle. The engine burns two fuels (hydrogen and a hydrocarbon) with oxygen in separate combustors. Combustion gases expand on a linear aerospike nozzle. An engine preliminary design is discussed. Attention is given to the evaluation process for selecting the optimum number of modules or divisions of the engine, aspects of cooling and power cycle balance, and details of engine operation.

Saturn V S-IC (First) Stage

NASA Technical Reports Server (NTRS)

2004-01-01

This cutaway illustration shows the Saturn V S-IC (first) stage with detailed callouts of the components. The S-IC Stage is 138 feet long and 33 feet in diameter, producing 7,500,000 pounds of thrust through five F-1 engines that are powered by liquid oxygen and kerosene. Four of the engines are mounted on an outer ring and gimbal for control purposes. The fifth engine is rigidly mounted in the center. When ignited, the roar produced by the five engines equals the sound of 8,000,000 hi-fi sets.

Saturn V Second Stage (S-II) Ready for Static Test

NASA Technical Reports Server (NTRS)

1965-01-01

Two workers are dwarfed by the five J-2 engines of the Saturn V second stage (S-II) as they make final inspections prior to a static test firing by North American Space Division. These five hydrogen -fueled engines produced one million pounds of thrust, and placed the Apollo spacecraft into earth orbit before departing for the moon. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Numerical analyses of a rocket engine turbine and comparison with air test data

NASA Technical Reports Server (NTRS)

Tran, Ken; Chan, Daniel C.; Hudson, Susan T.; Gaddis, Stephen W.

1992-01-01

The study presents cold air test data on the Space Shuttle Main Engine High Pressure Fuel Turbopump turbine recently collected at the NASA Marshall Space Flight Center. Overall performance data, static pressures on the first- and second-stage nozzles, and static pressures along with the gas path at the hub and tip are gathered and compared with various (1D, quasi-3D, and 3D viscous) analysis procedures. The results of each level of analysis are compared to test data to demonstrate the range of applicability for each step in the design process of a turbine. One-dimensional performance prediction, quasi-3D loading prediction, 3D wall pressure distribution prediction, and 3D viscous wall pressure distribution prediction are illustrated.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

The business end of a Second Stage (S-II) slowly emerges from the shipping container as workers prepare to transport the Saturn V component to the testing facility at MSFC. The Second Stage (S-II) underwent vibration and engine firing tests. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

First Stage Acceptance Test

NASA Technical Reports Server (NTRS)

1960-01-01

This photograph shows the intense smoke and fire created by the five F-1 engines from a test firing of the Saturn V first stage (S-1C) in the S-1C test stand at the Marshall Space Flight Center. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Aft View of Saturn V Third Stage (S-IVB)

NASA Technical Reports Server (NTRS)

1960-01-01

The powerful J-2 engine is prominent in this photograph of a Saturn V Third Stage (S-IVB) resting on a transporter in the Manufacturing Facility at Marshall Space Flight Center in Huntsville, Alabama. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

40 CFR 94.215 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2010 CFR

2010-07-01

... testing, to translation of designs from the test stage to the production stage, or to engine manufacture... certification under this subpart including: (A) In the case where a current production engine is modified for... production engine, a general description of the buildup of the engine (e.g., whether experimental heads were...

78 FR 28161 - Airworthiness Directives; Rolls-Royce plc Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2013-05-14

... Turbofan Engines AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Notice of proposed rulemaking... turbofan engines that have a high-pressure (HP) compressor stage 1 to 4 rotor disc installed, with a..., -524H-T-36, and -524H2-T-19 turbofan engines that have a HP compressor stage 1 to 4 rotor disc installed...

Main Chamber and Preburner Injector Technology

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Merkle, Charles L.

1999-01-01

This document reports the experimental and analytical research carried out at the Penn State Propulsion Engineering Research Center in support of NASA's plan to develop advanced technologies for future single stage to orbit (SSTO) propulsion systems. The focus of the work is on understanding specific technical issues related to bi-propellant and tri-propellant thrusters. The experiments concentrate on both cold flow demonstrations and hot-fire uni-element tests to demonstrate concepts that can be incorporated into hardware design and development. The analysis is CFD-based and is intended to support the design and interpretation of the experiments and to extrapolate findings to full-scale designs. The research is divided into five main categories that impact various SSTO development scenarios. The first category focuses on RP-1/gaseous hydrogen (GH2)/gaseous oxygen (GO2) tri-propellant combustion with specific emphasis on understanding the benefits of hydrogen addition to RP-1/oxygen combustion and in developing innovative injector technology. The second category investigates liquid oxygen (LOX)/GH2 combustion at main chamber near stoichiometric conditions to improve understanding of existing LOX/GH2 rocket systems. The third and fourth categories investigate the technical issues related with oxidizer-rich and fuel-rich propulsive concepts, issues that are necessary for developing the full-flow engine cycle. Here, injector technology issues for both LOX/GH2 and LOX/RP-1 propellants are examined. The last category, also related to the full-flow engine cycle, examines injector technology needs for GO2/GH2 propellant combustion at near-stoichiometric conditions for main chamber application.

Fiber-reinforced ceramic composites for Earth-to-orbit rocket engine turbines

NASA Technical Reports Server (NTRS)

Brockmeyer, Jerry W.; Schnittgrund, Gary D.

1990-01-01

Fiber reinforced ceramic matrix composites (FRCMC) are emerging materials systems that offer potential for use in liquid rocket engines. Advantages of these materials in rocket engine turbomachinery include performance gain due to higher turbine inlet temperature, reduced launch costs, reduced maintenance with associated cost benefits, and reduced weight. This program was initiated to assess the state of FRCMC development and to propose a plan for their implementation into liquid rocket engine turbomachinery. A complete range of FRCMC materials was investigated relative to their development status and feasibility for use in the hot gas path of earth-to-orbit rocket engine turbomachinery. Of the candidate systems, carbon fiber-reinforced silicon carbide (C/SiC) offers the greatest near-term potential. Critical hot gas path components were identified, and the first stage inlet nozzle and turbine rotor of the fuel turbopump for the liquid oxygen/hydrogen Space Transportation Main Engine (STME) were selected for conceptual design and analysis. The critical issues associated with the use of FRCMC were identified. Turbine blades were designed, analyzed and fabricated. The Technology Development Plan, completed as Task 5 of this program, provides a course of action for resolution of these issues.

Combustion Stability Verification for the Thrust Chamber Assembly of J-2X Developmental Engines 10001, 10002, and 10003

NASA Technical Reports Server (NTRS)

Morgan, C. J.; Hulka, J. R.; Casiano, M. J.; Kenny, R. J.; Hinerman, T. D.; Scholten, N.

2015-01-01

The J-2X engine, a liquid oxygen/liquid hydrogen propellant rocket engine available for future use on the upper stage of the Space Launch System vehicle, has completed testing of three developmental engines at NASA Stennis Space Center. Twenty-one tests of engine E10001 were conducted from June 2011 through September 2012, thirteen tests of the engine E10002 were conducted from February 2013 through September 2013, and twelve tests of engine E10003 were conducted from November 2013 to April 2014. Verification of combustion stability of the thrust chamber assembly was conducted by perturbing each of the three developmental engines. The primary mechanism for combustion stability verification was examining the response caused by an artificial perturbation (bomb) in the main combustion chamber, i.e., dynamic combustion stability rating. No dynamic instabilities were observed in the TCA, although a few conditions were not bombed. Additional requirements, included to guard against spontaneous instability or rough combustion, were also investigated. Under certain conditions, discrete responses were observed in the dynamic pressure data. The discrete responses were of low amplitude and posed minimal risk to safe engine operability. Rough combustion analyses showed that all three engines met requirements for broad-banded frequency oscillations. Start and shutdown transient chug oscillations were also examined to assess the overall stability characteristics, with no major issues observed.

NASA/DOD Aerospace Knowledge Diffusion Research Project. Paper 53: From student to entry-level professional: Examining the technical communications practices of early career-stage US aerospace engineers and scientists

NASA Technical Reports Server (NTRS)

Pinelli, Thomas E.; Holloway, Karen; Barclay, Rebecca O.; Kennedy, John M.

1995-01-01

Studies indicate that communications and information-related activities take up a substantial portion of an engineer's work week; therefore, effective communications and information-use skills are one of the key engineering competencies that early career-stage aerospace engineers and scientists must possess to be successful. Feedback from industry rates communications and information-use skills high in terms of their importance to engineering practice; however, this same feedback rates the communications and information-use skills of early career-stage engineers low. To gather adequate and generalizable data about the communications and information-related activities of entry-level aerospace engineers and scientists, we surveyed 264 members of the AIAA who have no more than 1-5 years of aerospace engineering work experience. To learn more about the concomitant communications norms, we compared the results of this study with data (1,673 responses) we collected from student members of the AIAA and with data (341 responses) we collected from a study of aerospace engineering professionals. In this paper, we report selected results from these studies that focused on the communications practices and information-related activities of early career-stage U.S. aerospace engineers and scientists in the workplace.

Orbit transfer vehicle engine study. Volume 2: Technical report

NASA Technical Reports Server (NTRS)

1980-01-01

The orbit transfer vehicle (OTV) engine study provided parametric performance, engine programmatic, and cost data on the complete propulsive spectrum that is available for a variety of high energy, space maneuvering missions. Candidate OTV engines from the near term RL 10 (and its derivatives) to advanced high performance expander and staged combustion cycle engines were examined. The RL 10/RL 10 derivative performance, cost and schedule data were updated and provisions defined which would be necessary to accommodate extended low thrust operation. Parametric performance, weight, envelope, and cost data were generated for advanced expander and staged combustion OTV engine concepts. A prepoint design study was conducted to optimize thrust chamber geometry and cooling, engine cycle variations, and controls for an advanced expander engine. Operation at low thrust was defined for the advanced expander engine and the feasibility and design impact of kitting was investigated. An analysis of crew safety and mission reliability was conducted for both the staged combustion and advanced expander OTV engine candidates.

Decision problems in management of construction projects

NASA Astrophysics Data System (ADS)

Szafranko, E.

2017-10-01

In a construction business, one must oftentimes make decisions during all stages of a building process, from planning a new construction project through its execution to the stage of using a ready structure. As a rule, the decision making process is made more complicated due to certain conditions specific for civil engineering. With such diverse decision situations, it is recommended to apply various decision making support methods. Both, literature and hands-on experience suggest several methods based on analytical and computational procedures, some less and some more complex. This article presents the methods which can be helpful in supporting decision making processes in the management of civil engineering projects. These are multi-criteria methods, such as MCE, AHP or indicator methods. Because the methods have different advantages and disadvantages, whereas decision situations have their own specific nature, a brief summary of the methods alongside some recommendations regarding their practical applications has been given at the end of the paper. The main aim of this article is to review the methods of decision support and their analysis for possible use in the construction industry.

Propulsion Ground Testing with High Test Peroxide: Lessons Learned

NASA Technical Reports Server (NTRS)

Bruce, Robert; Taylor, Gary; Taliancich, Paula

2002-01-01

Propulsion Ground Testing with High Test Peroxide (85 to 98% concentration) began at the NASA John C. Stennis Space Center in calendar year 1998, when the E3 Test Facility was modified to accomodate hydrogen peroxide (H2O2) in order to suport the research and development testing of the USAF Upper Stage Flight Experiment rocket engine. Since that time, efforts have continued to provide actual and planned test services to various customers, both U.S. Government and Commercial, in the ground test of many test articles, ranging from gas generators, to catalyst beds, to turbomachinery, to main injectors, to combustion chambers, to integrated rocket engines, to integrated stages. Along this path, and over the past 4 years, there has been both the rediscovery of previously learned lessons, through literature search, archive review, and personal interviews, as well as the learning of many new lessons as new areas are explored and new endeavors are tried. This paper will summarize those lessons learned in an effort to broaden the knowledge base as High Test Peroxide is considered more widely for use in rocket propulsion applications.

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.

Saturn V First Stage (S-1C) At MSFC

NASA Technical Reports Server (NTRS)

1960-01-01

This small group of unidentified officials is dwarfed by the gigantic size of the Saturn V first stage (S-1C) at the shipping area of the Manufacturing Engineering Laboratory at Marshall Space Flight Center in Huntsville, Alabama. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This cutaway illustration shows the Saturn V S-II (second) stage with callouts of major components. When the Saturn V first stage burns out and drops away, power for the Saturn was provided by the S-II (second) stage with five J-2 engines which produced a total of 1,150,000 pounds of thrust. Four outer engines are placed in a square pattern with gimbaling capability for control and guidance, with the fifth engine fixed rigidly in the center.

76 FR 76027 - Airworthiness Directives; Pratt & Whitney Division (PW) PW4000 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2011-12-06

... performed by PW. This AD requires removing certain part number (P/N) high-pressure turbine (HPT) stage 1 and... engines, with high-pressure turbine (HPT) stage 1 airseal, part number (P/N) 50L879; HPT stage 2 airseal...

Saturn Apollo Program

NASA Image and Video Library

1963-12-05

The test laboratory of the Marshall Space Flight Center (MSFC) tested the F-1 engine, the most powerful rocket engine ever fired at MSFC. The engine was tested on the newly modified Saturn IB Static Test Stand which had been used for three years to test the Saturn I eight-engine booster, S-I (first) stage. In 1961 the test stand was modified to permit static firing of the S-I/S-IB stage and the name of the stand was then changed to the S-IB Static Test Stand. Producing a combined thrust of 7,500,000 pounds, five F-1 engines powered the S-IC (first) stage of the Saturn V vehicle for the marned lunar mission.

Saturn Apollo Program

NASA Image and Video Library

1963-12-01

The test laboratory of the Marshall Space Flight Center (MSFC) tested the F-1 engine, the most powerful rocket engine ever fired at MSFC. The engine was tested on the newly modified Saturn IB static test stand that had been used for three years to test the Saturn I eight-engine booster, S-I (first) stage. In 1961, the test stand was modified to permit static firing of the S-I/S-IB stage and the name of the stand was then changed to the S-IB Static Test Stand. Producing a combined thrust of 7,500,000 pounds, five F-1 engines powered the S-IC (first) stage of the Saturn V vehicle for the marned lunar mission.

NASA's SPACE LAUNCH SYSTEM: Development and Progress

NASA Technical Reports Server (NTRS)

Honeycutt, John; Lyles, Garry

2016-01-01

NASA is embarked on a new era of space exploration that will lead to new capabilities, new destinations, and new discoveries by both human and robotic explorers. Today, the International Space Station (ISS) and robotic probes are yielding knowledge that will help make this exploration possible. NASA is developing both the Orion crew vehicle and the Space Launch System (SLS) (Figure 1), that will carry out a series of increasingly challenging missions leading to human exploration of Mars. This paper will discuss the development and progress on the SLS. The SLS architecture was designed to be safe, affordable, and sustainable. The current configuration is the result of literally thousands of trade studies involving cost, performance, mission requirements, and other metrics. The initial configuration of SLS, designated Block 1, will launch a minimum of 70 metric tons (mT) (154,324 pounds) into low Earth orbit - significantly greater capability than any current launch vehicle. It is designed to evolve to a capability of 130 mT (286,601 pounds) through the use of upgraded main engines, advanced boosters, and a new upper stage. With more payload mass and volume capability than any existing rocket, SLS offers mission planners larger payloads, faster trip times, simpler design, shorter design cycles, and greater opportunity for mission success. Since the program was officially created in fall 2011, it has made significant progress toward launch readiness in 2018. Every major element of SLS continued to make significant progress in 2015. Engineers fired Qualification Motor 1 (QM-1) in March 2015 to test the 5-segment motor, including new insulation, joint, and propellant grain designs. More than 70 major components of test article and flight hardware for the Core Stage have been manufactured. Seven test firings have been completed with an RS-25 engine under SLS operating conditions. The test article for the Interim Cryogenic Propulsion Stage (ICPS) has also been completed. Major work continues in 2016 as the program continues both flight and development RS-25 engine testing, begins welding test article and flight core stage tanks, completes stage adapter manufacturing, and test fires the second booster qualification motor. This paper will discuss the program's key accomplishments to date and the challenging work ahead for what will be the world's most capable launch vehicle.

Ram booster

NASA Technical Reports Server (NTRS)

Brand, Vance D. (Inventor); Morgan, Walter Ray (Inventor)

2011-01-01

The present invention is a space launch system and method to propel a payload bearing craft into earth orbit. The invention has two, or preferably, three stages. The upper stage has rocket engines capable of carrying a payload to orbit and provides the capability of releasably attaching to the lower, or preferably, middle stage. Similar to the lower stage, the middle stage is a reusable booster stage that employs all air breathing engines, is recoverable, and can be turned-around in a short time between missions.

Saturn Apollo Program

NASA Image and Video Library

1967-03-01

The Saturn V configuration is shown in inches and meters as illustrated by the Boeing Company. The Saturn V vehicle consisted of three stages: the S-IC (first) stage powered by five F-1 engines, the S-II (second) stage powered by five J-2 engines, the S-IVB (third) stage powered by one J-2 engine. A top for the first three stages was designed to contain the instrument unit, the guidance system, the Apollo spacecraft, and the escape system. The Apollo spacecraft consisted of the lunar module, the service module, and the command module. The Saturn V was designed perform lunar and planetary missions and it was capable of placing 280,000 pounds into Earth orbit.

A semireusable launch vehicle concept as a reference system for reusability analyses

NASA Astrophysics Data System (ADS)

Kleinau, W.

A two-stage concept called AR-X1, which uses H2O2 propellant and the HM 60 engine is presented. The first stage is reusable, the second expendable. The use of LH2/LOX in the first stage reduces the number of stages for geosynchronous transfer orbit (GTO) missions because of the higher performance. An 8 Mg payload can be injected in GTO (launch mass = 435 Mg). The first stage comprises four parallel stretched second stage tanks with 320 Mg propellants (total) and eight HM 60 engines arranged within the heat shield, plus one central HM 60 thruster for the soft landing maneuver. Engine performance is increased by adapting the expansion ratio to the external pressure. Trajectory calculations show that the first stage flight range is 1 500 km. Braking before touchdown is performed by retro thrust, requiring 2.5 to 3 Mg propellants. First-stage reuse reduces cost per launch by 50% compared with an expendable three stage design.

Optimization, an Important Stage of Engineering Design

ERIC Educational Resources Information Center

Kelley, Todd R.

2010-01-01

A number of leaders in technology education have indicated that a major difference between the technological design process and the engineering design process is analysis and optimization. The analysis stage of the engineering design process is when mathematical models and scientific principles are employed to help the designer predict design…

40 CFR 89.2 - Definitions.

Code of Federal Regulations, 2011 CFR

2011-07-01

... secondarily for operation in water. Auxiliary emission control device (AECD) means any element of design that... design which controls or reduces the emission of substances from an engine. Engine, as used in this part... testing, to translation of designs from the test stage to the production stage, or to engine manufacture...

40 CFR 89.2 - Definitions.

Code of Federal Regulations, 2012 CFR

2012-07-01

... secondarily for operation in water. Auxiliary emission control device (AECD) means any element of design that... design which controls or reduces the emission of substances from an engine. Engine, as used in this part... testing, to translation of designs from the test stage to the production stage, or to engine manufacture...

40 CFR 89.2 - Definitions.

Code of Federal Regulations, 2014 CFR

2014-07-01

... secondarily for operation in water. Auxiliary emission control device (AECD) means any element of design that... design which controls or reduces the emission of substances from an engine. Engine, as used in this part... testing, to translation of designs from the test stage to the production stage, or to engine manufacture...

78 FR 47581 - Airworthiness Directives; Turbomeca S.A. Turboshaft Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-06

... turbine disc. We are proposing this AD to prevent disc cracking, uncontained 2nd-stage turbine blade..., possibly resulting in uncontained second stage turbine blade release with consequent damage to, and reduced...-stage turbine blade release, damage to the engine, and damage to the helicopter. (e) Actions and...

Second stage of Saturn V being assembled with the first stage.

NASA Technical Reports Server (NTRS)

1965-01-01

The hydrogen-powered second stage is being lowered into place during the final phase of fabrication of the Saturn V moon rocket at North American's Seal Beach, California facility. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Developmental Flight Instrumentation System for the Crew Launch Vehicle

NASA Technical Reports Server (NTRS)

Crawford, Kevin; Thomas, John

2006-01-01

The National Aeronautics and Space Administration is developing a new launch vehicle to replace the Space Shuttle. The Crew Launch Vehicle (CLV) will be a combination of new design hardware and heritage Apollo and Space Shuttle hardware. The current CLV configuration is a 5 segment solid rocket booster first stage and a new upper stage design with a modified Apollo era J-2 engine. The current schedule has two test flights with a first stage and a structurally identical, but without engine, upper stage. Then there will be two more test flights with a full complement of flight hardware. After the completion of the test flights, the first manned flight to the International Space Station is scheduled for late 2012. To verify the CLV's design margins a developmental flight instrumentation (DFI) system is needed. The DFI system will collect environmental and health data from the various CLV subsystem's and either transmit it to the ground or store it onboard for later evaluation on the ground. The CLV consists of 4 major elements: the first stage, the upper stage, the upper stage engine and the integration of the first stage, upper stage and upper stage engine. It is anticipated that each of CLVs elements will have some version of DFI. This paper will discuss a conceptual DFI design for each element and also of an integrated CLV DFI system.

Thrust vector control of upper stage with a gimbaled thruster during orbit transfer

NASA Astrophysics Data System (ADS)

Wang, Zhaohui; Jia, Yinghong; Jin, Lei; Duan, Jiajia

2016-10-01

In launching Multi-Satellite with One-Vehicle, the main thruster provided by the upper stage is mounted on a two-axis gimbal. During orbit transfer, the thrust vector of this gimbaled thruster (GT) should theoretically pass through the mass center of the upper stage and align with the command direction to provide orbit transfer impetus. However, it is hard to be implemented from the viewpoint of the engineering mission. The deviations of the thrust vector from the command direction would result in large velocity errors. Moreover, the deviations of the thrust vector from the upper stage mass center would produce large disturbance torques. This paper discusses the thrust vector control (TVC) of the upper stage during its orbit transfer. Firstly, the accurate nonlinear coupled kinematic and dynamic equations of the upper stage body, the two-axis gimbal and the GT are derived by taking the upper stage as a multi-body system. Then, a thrust vector control system consisting of the special attitude control of the upper stage and the gimbal rotation of the gimbaled thruster is proposed. The special attitude control defined by the desired attitude that draws the thrust vector to align with the command direction when the gimbal control makes the thrust vector passes through the upper stage mass center. Finally, the validity of the proposed method is verified through numerical simulations.

Orbital Transfer Vehicle (OTV) engine study. Phase A: Extension

NASA Technical Reports Server (NTRS)

Sobin, A. J.

1980-01-01

The current Phase A-Extension of the OTV engine study program aims to provide additional expander and staged combustion cycle data that will lead to design definition of the OTV engine. The proposed program effort seeks to optimize the expander cycle engine concept (consistent with identified OTV engine requirements), investigate the feasibility of kitting the staged combustion cycle engine to provide extended thrust operation, and conduct in-depth analysis of development risk, crew safety, and reliability for both cycles. Additional tasks address the costing of a 10/K thrust expander cycle engine and support of OTV systems study contractors.

NACA Engineer Examines Wind Tunnel Compressor Blades

NASA Image and Video Library

1955-09-21

An engineer examines the main compressor for the 10- by 10-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The engineers were preparing the new wind tunnel for its initial runs in early 1956. The 10- by 10 was the most powerful propulsion wind tunnel in the nation. The facility was part of Congress’ Unitary Plan Act which coordinated wind tunnel construction at the NACA, Air Force, industry, and universities. The 10- by 10 was the largest of the three NACA tunnels built under the act. The 20-foot diameter eight-stage axial flow compressor, seen in this photograph, could generate air flows up to Mach 2.5 through the test section. The stainless steel compressor had 584 blades ranging from 1.8 to 3.25 feet in length. This main compressor was complemented by a secondary axial flow compressor. Working in tandem the two could generate wind streams up to Mach 3.5. The Cleveland Chamber of Commerce presented NACA Lewis photographer Bill Bowles with a second place award for this photograph in their Business and Professional category. The photograph was published in October 1955 edition of its periodical, The Clevelander, which highlighted local professional photographers. Fellow Lewis photographer Gene Giczy won second place in another category for a photograph of Cleveland Municipal Airport.

Undergraduate environmental engineering education in China

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

Yang, C.; Bero, B.N.

1999-07-01

In this paper, the development process, present situations, causes of improvement, and trends of higher education of environmental engineering in China are discussed. Several education modes in environmental engineering in China are also presented. The development process can be divided into three stages: the beginning stage, the expansion stage, and the modification stage. The 1970's and early 1980's wake of environmental consciousness and serious pollution situation in China resulted in about 20 universities setting up an environmental engineering specialty. The late 1980's and middle 1990's job opportunities for undergraduates in China resulted in many universities' creation of the environmental engineeringmore » specialty from specialties such as geography, geology, hydrology, mining engineering, and mineral separation engineering where job opportunities were stagnant. At present, adjustment and improvement of environmental engineering education are urgently required because of the excessive increase of undergraduate number, change of job opportunities and implementation of five-work-day system in China. Other problems include how to determine the ratio of social science courses to engineering science courses, how to determine the relationship of fundamental and applied courses, and how to determine the specialized direction. Hunan University, as a typical university conferring an accredited Bachelor degree in Environmental Engineering in four academic years in China, has been improving the instruction schedule for undergraduate education in environmental engineering. The curricula of the three phases for undergraduates of environmental engineering specialty at Hunan University are presented as a case study.« less

Enhancement of utilization of encryption engine

DOEpatents

Robertson, Robert J.; Witzke, Edward L.

2008-04-22

A method of enhancing throughput of a pipelined encryption/decryption engine for an encryption/decryption process has a predetermined number of stages and provides feedback around the stages (and of such an encryption/decryption engine) by receiving a source datablock for a given stage and encryption/decryption context identifier; indexing according to the encryption/decryption context identifier into a bank of initial variables to retrieve an initial variable for the source datablock; and generating an output datablock from the source datablock and its corresponding initial variable.

Supersonic Transport Noise Reduction Technology Program - Phase 2, Volume 2

DTIC Science & Technology

1975-09-01

a J85 is shown on Figure 350. The J85 turbojet engine has an eight-stage compressor (with an air bleed system) and a two-stage turbine . Blade ...investigated in this program using a YJ85 engine . Both turbine second-stage spacing ( blade - vane ) and exhaust duct treatment were determined to be...using a J85 engine with massive Inlet suppressor and open nozzle to unmask the turbine . Second-stag« turbine blade /nozzle spacing and exhaust

Saturn V First Stage Lowered to the Ground After Static Test

NASA Technical Reports Server (NTRS)

1966-01-01

This vintage photograph shows the 138-foot long first stage of the Saturn V being lowered to the ground following a successful static test firing at Marshall Space flight Center's S-1C test stand. The firing provided NASA engineers information on the booster's systems. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

H-1 engine characteristics: The H-1 engine was developed under the management of the Marshall Space Flight Center (MSFC). The cluster of eight H-1 engines was used to power the first stage of the Saturn I (S-I stage) and Saturn IB (S-IVB stage) launch vehicles, and produced 188,00 pounds of thrust, a combined thrust of 1,500,000 pounds, later uprated to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program.

Conceptual design of a two-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

1991-01-01

A conceptual design study of a two-stage-to-orbit vehicle is presented. Three configurations were initially investigated with one configuration selected for further development. The major objective was to place a 20,000-lb payload into a low Earth orbit using a two-stage vehicle. The first stage used air-breathing engines and employed a horizontal takeoff, while the second stage used rocket engines to achieve a 250-n.m. orbit. A two-stage-to-orbit vehicle seems a viable option for the next-generation space shuttle.

KSC-05PD-1592

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Media gather in the television studio at the NASA News Center to hear members of the Mission Management Team reveal aspects of the troubleshooting and testing being done on the liquid hydrogen tank low-level fuel cut-off sensor. On the stage at right are (from left) Wayne Hale, Space Shuttle deputy program manager; John Muratore, manager of Systems Engineering and Integration for the Space Shuttle Program; and Mike Wetmore, director of Space Shuttle Processing. The sensor failed a routine prelaunch check during the launch countdown July 13, causing mission managers to scrub Discovery's first launch attempt. The sensor protects the Shuttle's main engines by triggering their shutdown in the event fuel runs unexpectedly low. The sensor is one of four inside the liquid hydrogen section of the External Tank (ET).

Using Compilers to Enhance Cryptographic Product Development

NASA Astrophysics Data System (ADS)

Bangerter, E.; Barbosa, M.; Bernstein, D.; Damgård, I.; Page, D.; Pagter, J. I.; Sadeghi, A.-R.; Sovio, S.

Developing high-quality software is hard in the general case, and it is significantly more challenging in the case of cryptographic software. A high degree of new skill and understanding must be learnt and applied without error to avoid vulnerability and inefficiency. This is often beyond the financial, manpower or intellectual resources avail-able. In this paper we present the motivation for the European funded CACE (Computer Aided Cryptography Engineering) project The main objective of CACE is to provide engineers (with limited or no expertise in cryptography) with a toolbox that allows them to generate robust and efficient implementations of cryptographic primitives. We also present some preliminary results already obtained in the early stages of this project, and discuss the relevance of the project as perceived by stakeholders in the mobile device arena.

Ares V and RS-68B

NASA Technical Reports Server (NTRS)

Creech, Steve; Taylor, Jim; Bellamy, Scott; Kuck, Fritz

2008-01-01

Ares V is the heavy lift vehicle NASA is designing for lunar and other space missions. It has significantly more lift capability than the Saturn V vehicle used for the Apollo missions to the moon. Ares V is powered by two recoverable 5.5 segment solid rocket boosters and six RS-68B engines on the core stage. The upper stage, designated as the Earth Departure Stage, is powered by a single J-2X engine. This paper provides an overview of the Ares V vehicle and the RS-68B engine, an upgrade to the Pratt & Whitney Rocketdyne RS-68 engine developed for the Delta IV vehicle.

Saturn Apollo Program

NASA Image and Video Library

1961-01-01

The Saturn I S-I stage with eight H-1 engines, located in Marshall Space Flight Center building 4705, showing the positioning of eight H-1 engines. The Saturn I S-I stage had eight H-1 engines clustered, using liquid oxygen/kerosene-1 (LOX/RP-1) propellants capable of producing a total of 1,500,000 pounds of thrust.

Engine-Out Capabilities Assessment of Heavy Lift Launch Vehicles

NASA Technical Reports Server (NTRS)

Holladay, Jon; Baggett, Keithe; Thrasher, Chad; Bellamy, K. Scott; Feldman, Stuart

2012-01-01

Engine-out (EO) is a condition that might occur during flight due to the failure of one or more engines. Protection against this occurrence can be called engine-out capability (EOC) whereupon significantly improved loss of mission may occur, in addition to reduction in performance and increased cost. A standardized engine-out capability has not been studied exhaustively as it pertains to space launch systems. This work presents results for a specific vehicle design with specific engines, but also uniquely provides an approach to realizing the necessity of EOC for any launch vehicle system design. A derived top-level approach to engine-out philosophy for a heavy lift launch vehicle is given herein, based on an historical assessment of launch vehicle capabilities. The methodology itself is not intended to present a best path forward, but instead provides three parameters for assessment of a particular vehicle. Of the several parameters affected by this EOC, the three parameters of interest in this research are reliability (Loss of Mission (LOM) and Loss of Crew (LOC)), vehicle performance, and cost. The intent of this effort is to provide insight into the impacts of EO capability on these parameters. The effects of EOC on reliability, performance and cost are detailed, including how these important launch vehicle metrics can be combined to assess what could be considered overall launch vehicle affordability. In support of achieving the first critical milestone (Mission Concept Review) in the development of the Space Launch System (SLS), a team assessed two-stage, large-diameter vehicles that utilized liquid oxygen (LOX)-RP propellants in the First Stage and LOX/LH2 propellant in the Upper Stage. With multiple large thrust-class engines employed on the stages, engine-out capability could be a significant driver to mission success. It was determined that LOM results improve by a factor of five when assuming EOC for both Core Stage (CS) (first stage) and Upper Stage (US) EO, assuming a reference launch vehicle with 5 RP engines on the CS and 3 LOX/LH2 engines on the US. The benefit of adding both CS and US engine-out capability is significant. When adding EOC for either first or second stages, there is less than a 20% benefit. Performance analysis has shown that if the vehicle is not protected for EO during the first part of the flight and only protected in the later part of the flight, there is a diminishing performance penalty, as indicated by failures occurring in the first stage at different times. This work did not consider any options to abort. While adding an engine for EOC drives cost upward, the impact depends on the number of needed engines manufactured per year and the launch manifest. There is a significant cost savings if multiple flights occur within one year. Flying two flights per year would cost approximately $4,000 per pound less than the same configuration with one flight per year, assuming both CS and US EOC. The cost is within 15% of the cost of one flight per year with no engine-out capability for the same vehicle. This study can be extended to other launch vehicles. While the numbers given in this paper are specific to a certain vehicle configuration, the process requires only a high level of data to allow an analyst to draw conclusions. The weighting of each of the identified parameters will determine the optimization of each launch vehicle. The results of this engine-out assessment provide a means to understand this optimization while maintaining an unbiased perspective.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This is a cutaway illustration of the Saturn V launch vehicle with callouts of the major components. The Saturn V is the largest and most powerful launch vehicle developed in the United States. It was a three stage rocket, 363 feet in height, used for sending American astronauts to the moon and for placing the Skylab in Earth orbit. The Saturn V was designed to perform Earth orbital missions through the use of the first two stages, while all three stages were used for lunar expeditions. The S-IC stage (first stage) was powered by five F- engines, which burned kerosene and liquid oxygen to produce more than 7,500,000 pounds of thrust. The S-II (second) stage was powered by five J-2 engines, that burned liquid hydrogen and liquid oxygen and produced 1,150,000 pounds thrust. The S-IVB (third) stage used one J-2 engine, producing 230,000 pounds of thrust, with a re-start capability. The Marshall Space Flight Center and its contractors designed, developed, and assembled the Saturn V launch vehicle stages.

[Insights into engineering of cellulosic ethanol].

PubMed

Yue, Guojun; Wu, Guoqing; Lin, Xin

2014-06-01

For energy security, air pollution concerns, coupled with the desire to sustain the agricultural sector and revitalize the rural economy, many countries have applied ethanol as oxygenate or fuel to supplement or replace gasoline in transportation sector. Because of abundant feedstock resources and effective reduction of green-house-gas emissions, the cellulosic ethanol has attracted great attention. With a couple of pioneers beginning to produce this biofuel from biomass in commercial quantities around the world, it is necessary to solve engineering problems and complete the economic assessment in 2015-2016, gradually enter the commercialization stage. To avoid "competing for food with humans and competing for land with food", the 1st generation fuel ethanol will gradually transit to the 2nd generation cellulosic ethanol. Based on the overview of cellulosic ethanol industrialization from domestic and abroad in recent years, the main engineering application problems encountered in pretreatment, enzymes and enzymatic hydrolysis, pentose/hexose co-fermentation strains and processes, equipment were discussed from chemical engineering and biotechnology perspective. The development direction of cellulosic ethanol technology in China was addressed.

Fabrication and Testing of Low Cost 2D Carbon-Carbon Nozzle Extensions at NASA/MSFC

NASA Technical Reports Server (NTRS)

Greene, Sandra Elam; Shigley, John K.; George, Russ; Roberts, Robert

2015-01-01

Subscale liquid engine tests were conducted at NASA/MSFC using a 1.2 Klbf engine with liquid oxygen (LOX) and gaseous hydrogen. Testing was performed for main-stage durations ranging from 10 to 160 seconds at a chamber pressure of 550 psia and a mixture ratio of 5.7. Operating the engine in this manner demonstrated a new and affordable test capability for evaluating subscale nozzles by exposing them to long duration tests. A series of 2D C-C nozzle extensions were manufactured, oxidation protection applied and then tested on a liquid engine test facility at NASA/MSFC. The C-C nozzle extensions had oxidation protection applied using three very distinct methods with a wide range of costs and process times: SiC via Polymer Impregnation & Pyrolysis (PIP), Air Plasma Spray (APS) and Melt Infiltration. The tested extensions were about 6" long with an exit plane ID of about 6.6". The test results, material properties and performance of the 2D C-C extensions and attachment features will be discussed.

Third Stage (S-IVB) At KSC

NASA Technical Reports Server (NTRS)

1960-01-01

A NASA technician is dwarfed by the gigantic Third Stage (S-IVB) as it rests on supports in a facility at KSC. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4%. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

High-output LED-based light engine for profile lighting fixtures with high color uniformity using freeform reflectors.

PubMed

Gadegaard, Jesper; Jensen, Thøger Kari; Jørgensen, Dennis Thykjær; Kristensen, Peter Kjær; Søndergaard, Thomas; Pedersen, Thomas Garm; Pedersen, Kjeld

2016-02-20

In the stage lighting and entertainment market, light engines (LEs) for lighting fixtures are often based on high-intensity discharge (HID) bulbs. Switching to LED-based light engines gives possibilities for fast switching, additive color mixing, a longer lifetime, and potentially, more energy-efficient systems. The lumen output of a single LED is still not sufficient to replace an HID source in high-output profile fixtures, but combining multiple LEDs can create an LE with a similar output, but with added complexity. This paper presents the results of modeling and testing such a light engine. Custom ray-tracing software was used to design a high-output red, green and blue LED-based light engine with twelve CBT-90 LEDs using a dual-reflector principle. The simulated optical system efficiency was 0.626 with a perfect (R=1) reflector coating for light delivered on a target surface through the entire optical system. A profile lighting fixture prototype was created, and provided an output of 6744 lumen and an efficiency of 0.412. The lower efficiency was mainly due to a non-optimal reflector coating, and the optimized design is expected to reach a significantly higher efficiency.

Experimental clean combustor program, phase 2

NASA Technical Reports Server (NTRS)

Roberts, R.; Peduzzi, A.; Vitti, G. E.

1976-01-01

Combustor pollution reduction technology for commercial CTOL engines was generated and this technology was demonstrated in a full-scale JT9D engine in 1976. Component rig refinement of the two best combustor concepts were tested. These concepts are the vorbix combustor, and a hybrid combustor which combines the pilot zone of the staged premix combustor and the main zone of the swirl-can combustor. Both concepts significantly reduced all pollutant emissions relative to the JT9D-7 engine combustor. However, neither concept met all program goals. The hybrid combustor met pollution goals for unburned hydrocarbons and carbon monoxide but did not achieve the oxides of nitrogen goal. This combustor had significant performance deficiencies. The Vorbix combustor met goals for unburned hydrocarbons and oxides of nitrogen but did not achieve the carbon monoxide goal. Performance of the vorbix combustor approached the engine requirements. On the basis of these results, the vorbix combustor was selected for the engine demonstration program. A control study was conducted to establish fuel control requirements imposed by the low-emission combustor concepts and to identify conceptual control system designs. Concurrent efforts were also completed on two addendums: an alternate fuels addendum and a combustion noise addendum.

Modeling and simulation of a counter-rotating turbine system for underwater vehicles

NASA Astrophysics Data System (ADS)

Wang, Xinping; Dang, Jianjun

2016-12-01

The structure of a counter-rotating turbine of an underwater vehicle is designed by adding the counter-rotating second-stage turbine disk after the conventional single-stage turbine. The available kinetic energy and the absorption power of the auxiliary system are calculated at different working conditions, and the results show that the power of the main engine and auxiliary system at the counter-rotating turbine system matches well with each other. The experimental simulation of the lubricating oil loop, fuel loop, and seawater loop are completed right before the technology scheme of the counter-rotating turbine system is proposed. The simulation results indicate that the hydraulic transmission system can satisfy the requirements for an underwater vehicle running at a steady sailing or variable working conditions.

Calculation of Turbine Axial Thrust by Coupled CFD Simulations of the Main Flow Path and Secondary Cavity Flow in an SLI LOX Turbine

NASA Technical Reports Server (NTRS)

Dorney, D. J.; Marci, Bogdan; Tran, Ken; Sargent, Scott

2003-01-01

Each single reusable Space Launch Initiative (SLI) booster rocket is an engine operating at a record vacuum thrust level of over 730,000 Ibf using LOX and LH2. This thrust is more than 10% greater than that of the Delta IV rocket, resulting in relatively large LOX and LH2 turbopumps. Since the SLI rocket employs a staged combustion cycle the level of pressure is very high (thousands of psia). This high pressure creates many engineering challenges, including the balancing of axial-forces on the turbopumps. One of the main parameters in the calculation of the axial force is the cavity pressure upstream of the turbine disk. The flow in this cavity is very complex. The lack of understanding of this flow environment hinders the accurate prediction of axial thrust. In order to narrow down the uncertainty band around the actual turbine axial force, a coupled, unsteady computational methodology has been developed to simulate the interaction between the turbine main flow path and the cavity flow. The CORSAIR solver, an unsteady three- dimensional Navier-Stokes code for turbomachinery applications, was used to solve for both the main and the secondary flow fields. Turbine axial thrust values are presented in conjunction with the CFD simulation, together with several considerations regarding the turbine instrumentation for axial thrust estimations during test.

KSC-2014-3615

NASA Image and Video Library

2014-08-20

VANDENBERG AIR FORCE BASE, Calif. – The second stage of the Delta II rocket for NASA's Soil Moisture Active Passive mission, or SMAP, is transferred into the top of the mobile service tower at Space Launch Complex 2 on Vandenberg Air Force Base in California. Operations are underway to install the second stage atop the rocket's first stage. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch is scheduled for no earlier than November 2014. To learn more about SMAP, visit http://smap.jpl.nasa.gov. Photo credit: NASA/Randy Beaudoin

Differential Cloud Particles Evolution Algorithm Based on Data-Driven Mechanism for Applications of ANN

PubMed Central

2017-01-01

Computational scientists have designed many useful algorithms by exploring a biological process or imitating natural evolution. These algorithms can be used to solve engineering optimization problems. Inspired by the change of matter state, we proposed a novel optimization algorithm called differential cloud particles evolution algorithm based on data-driven mechanism (CPDD). In the proposed algorithm, the optimization process is divided into two stages, namely, fluid stage and solid stage. The algorithm carries out the strategy of integrating global exploration with local exploitation in fluid stage. Furthermore, local exploitation is carried out mainly in solid stage. The quality of the solution and the efficiency of the search are influenced greatly by the control parameters. Therefore, the data-driven mechanism is designed for obtaining better control parameters to ensure good performance on numerical benchmark problems. In order to verify the effectiveness of CPDD, numerical experiments are carried out on all the CEC2014 contest benchmark functions. Finally, two application problems of artificial neural network are examined. The experimental results show that CPDD is competitive with respect to other eight state-of-the-art intelligent optimization algorithms. PMID:28761438

Interstage Flammability Analysis Approach

NASA Technical Reports Server (NTRS)

Little, Jeffrey K.; Eppard, William M.

2011-01-01

The Interstage of the Ares I launch platform houses several key components which are on standby during First Stage operation: the Reaction Control System (ReCS), the Upper Stage (US) Thrust Vector Control (TVC) and the J-2X with the Main Propulsion System (MPS) propellant feed system. Therefore potentially dangerous leaks of propellants could develop. The Interstage leaks analysis addresses the concerns of localized mixing of hydrogen and oxygen gases to produce deflagration zones in the Interstage of the Ares I launch vehicle during First Stage operation. This report details the approach taken to accomplish the analysis. Specified leakage profiles and actual flammability results are not presented due to proprietary and security restrictions. The interior volume formed by the Interstage walls, bounding interfaces with the Upper and First Stages, and surrounding the J2-X engine was modeled using Loci-CHEM to assess the potential for flammable gas mixtures to develop during First Stage operations. The transient analysis included a derived flammability indicator based on mixture ratios to maintain achievable simulation times. Validation of results was based on a comparison to Interstage pressure profiles outlined in prior NASA studies. The approach proved useful in the bounding of flammability risk in supporting program hazard reviews.

Design and Analysis of a Turbopump for a Conceptual Expander Cycle Upper-Stage Engine

NASA Technical Reports Server (NTRS)

Dorney, Daniel J.; Rothermel, Jeffry; Griffin, Lisa W.; Thornton, Randall J.; Forbes, John C.; Skelly, Stephen E.; Huber, Frank W.

2006-01-01

As part of the development of technologies for rocket engines that will power spacecraft to the Moon and Mars, a program was initiated to develop a conceptual upper stage engine with wide flow range capability. The resulting expander cycle engine design employs a radial turbine to allow higher pump speeds and efficiencies. In this paper, the design and analysis of the pump section of the engine are discussed. One-dimensional meanline analyses and three-dimensional unsteady computational fluid dynamics simulations were performed for the pump stage. Configurations with both vaneless and vaned diffusers were investigated. Both the meanline analysis and computational predictions show that the pump will meet the performance objectives. Additional details describing the development of a water flow facility test are also presented.

SSME alternate turbopump (pump section) axial load analysis

NASA Technical Reports Server (NTRS)

Crease, G. A.; Rosello, A., Jr.; Fetfatsidis, A. K.

1989-01-01

A flow balancing computer program constructed to calculate the axial loads on the Space Shuttle Main Engine (SSME) alternate turbopumps (ATs) pump sections are described. The loads are used in turn to determine load balancing piston design requirements. The application of the program to the inlet section, inducer/impeller/stage, bearings, seals, labyrinth, damper, piston, face and corner, and stationary/rotating surfaces is indicated. Design analysis results are reported which show that the balancing piston's designs are adequate and that performance and life will not be degraded by the turbopump's axial load characteristics.

Centaur Engine Display Installation

NASA Image and Video Library

2016-04-14

The 6,600 pound Centaur test article is a rare artifact recently transported from the U.S. Space and Rocket Center in Alabama. Centaur, developed at NASA Glenn Research Center in the late 1950s, was the world's first high-energy upper stage, burning liquid hydrogen (LH2) and liquid oxygen (LOX), and has enabled the launch of some of NASA's most important scientific missions over its 50-year history. In this image, technicians prepare to mount the hardware on a permanent display stand close to the main entrance at NASA Glenn Research Center.

Hybridization of Architectural Styles for Integrated Enterprise Information Systems

NASA Astrophysics Data System (ADS)

Bagusyte, Lina; Lupeikiene, Audrone

Current enterprise systems engineering theory does not provide adequate support for the development of information systems on demand. To say more precisely, it is forming. This chapter proposes the main architectural decisions that underlie the design of integrated enterprise information systems. This chapter argues for the extending service-oriented architecture - for merging it with component-based paradigm at the design stage and using connectors of different architectural styles. The suitability of general-purpose language SysML for the modeling of integrated enterprise information systems architectures is described and arguments pros are presented.

40 CFR 86.091-7 - Maintenance of records; submittal of information; right of entry.

Code of Federal Regulations, 2010 CFR

2010-07-01

... certification testing, to translation of designs from the test stage to the production stage, or to vehicle (or...) In the case where a current production engine is modified for use in a certification vehicle (or as a... from a current production engine, a general description of the buildup of the engine (e.g...

Space Transportation Main Engine

NASA Technical Reports Server (NTRS)

Monk, Jan C.

1992-01-01

The topics are presented in viewgraph form and include the following: Space Transportation Main Engine (STME) definition, design philosophy, robust design, maximum design condition, casting vs. machined and welded forgings, operability considerations, high reliability design philosophy, engine reliability enhancement, low cost design philosophy, engine systems requirements, STME schematic, fuel turbopump, liquid oxygen turbopump, main injector, and gas generator. The major engine components of the STME and the Space Shuttle Main Engine are compared.

78 FR 48339 - Airworthiness Directives; Rolls-Royce Corporation Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-08

... currently requires removing certain high-pressure turbine (HPT) stage 2 wheels, or performing inspections on... turbofan engines: (1) With an installed high-pressure turbine (HPT) stage 2 wheel, part number (P/N...

77 FR 12755 - Airworthiness Directives; Pratt & Whitney (PW) Division Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-03-02

... pressure turbine (LPT) of certain PW4000-94'' and PW4000-100'' turbofan engines. These fractures caused an... stages. This AD requires reassembling the 2nd stage HPT blades at the next HPT overhaul and the 3rd stage LPT blades at the next LPT overhaul, using the latest assembly technique. The actions proposed in this...

CLOSEUP VIEW OF THE FIRST STAGE OF THE SATURN I ...

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

CLOSE-UP VIEW OF THE FIRST STAGE OF THE SATURN I ROCKET, SHOWING A DETAIL VIEW OF THE ENGINE CLUSTER. THE SATURN I ROCKET WAS THE FIRST UNITED STATES ROCKET TO HAVE MULTIPLE ENGINES ON A SINGLE STAGE. - Marshall Space Flight Center, Saturn Propulsion & Structural Test Facility, East Test Area, Huntsville, Madison County, AL

Two technicians apply insulation to S-II second stage

NASA Technical Reports Server (NTRS)

1964-01-01

Two technicians apply insulation to the outer surface of the S-II second stage booster for the Saturn V moon rocket. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

The Effect of Bypass Nozzle Exit Area on Fan Aerodynamic Performance and Noise in a Model Turbofan Simulator

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Podboy, Gary, G.; Woodward, Richard P.; Jeracki, Robert, J.

2013-01-01

The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted called the Fan Broadband Source Diagnostic Test as part of the NASA Quiet Aircraft Technology program. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator which represented a current generation, medium pressure ratio, high bypass turbofan aircraft engine. The investigation focused on simulating in model scale only the bypass section of the turbofan engine. The test objectives were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by determining their impact on the aerodynamic and acoustic performance of the fan stage; and conduct detailed flow diagnostics within the fan flow field to characterize the physics of the noise generation mechanisms in a turbofan model. This report discusses results obtained for one aspect of the Source Diagnostic Test that investigated the effect of the bypass or fan nozzle exit area on the bypass stage aerodynamic performance, specifically the fan and outlet guide vanes or stators, as well as the farfield acoustic noise level. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter flow diagnostic results are presented for the fan and four different fixed-area bypass nozzle configurations. The nozzles simulated fixed engine operating lines and encompassed the fan stage operating envelope from near stall to cruise. One nozzle was selected as a baseline reference, representing the nozzle area which would achieve the design point operating conditions and fan stage performance. The total area change from the smallest to the largest nozzle was 12.9 percent of the baseline nozzle area. The results will show that there are significant changes in aerodynamic performance and farfield acoustics as the fan nozzle area is increased. The weight flow through the fan model increased between 7 and 9 percent, the fan and stage pressure dropped between 8 and 10 percent, and the adiabatic efficiency increased between 2 and 3 percent--the magnitude of the change dependent on the fan speed. Results from force balance measurements of fan and outlet guide vane thrust will show that as the nozzle exit area is increased the combined thrust of the fan and outlet guide vanes together also increases, between 2 and 3.5 percent, mainly due to the increase in lift from the outlet guide vanes. In terms of farfield acoustics, the overall sound power level produced by the fan stage dropped nearly linearly between 1 dB at takeoff condition and 3.5 dB at approach condition, mainly due to a decrease in the broadband noise levels. Finally, fan swirl angle survey and Laser Doppler Velocimeter mean velocity and turbulence data obtained in the fan wake will show that the swirl angles and turbulence levels within the wake decrease as the fan nozzle area increases, which helps to explain the drop in the fan broadband noise at all fan speeds.

Atmospheric Mining in the Outer Solar System: Outer Planet Orbital Transfer and Lander Analyses

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2016-01-01

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 seconds. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

Human iPS cell-engineered cardiac tissue sheets with cardiomyocytes and vascular cells for cardiac regeneration

PubMed Central

Masumoto, Hidetoshi; Ikuno, Takeshi; Takeda, Masafumi; Fukushima, Hiroyuki; Marui, Akira; Katayama, Shiori; Shimizu, Tatsuya; Ikeda, Tadashi; Okano, Teruo; Sakata, Ryuzo; Yamashita, Jun K.

2014-01-01

To realize cardiac regeneration using human induced pluripotent stem cells (hiPSCs), strategies for cell preparation, tissue engineering and transplantation must be explored. Here we report a new protocol for the simultaneous induction of cardiomyocytes (CMs) and vascular cells [endothelial cells (ECs)/vascular mural cells (MCs)], and generate entirely hiPSC-engineered cardiovascular cell sheets, which showed advantageous therapeutic effects in infarcted hearts. The protocol adds to a previous differentiation protocol of CMs by using stage-specific supplementation of vascular endothelial cell growth factor for the additional induction of vascular cells. Using this cell sheet technology, we successfully generated physically integrated cardiac tissue sheets (hiPSC-CTSs). HiPSC-CTS transplantation to rat infarcted hearts significantly improved cardiac function. In addition to neovascularization, we confirmed that engrafted human cells mainly consisted of CMs in >40% of transplanted rats four weeks after transplantation. Thus, our HiPSC-CTSs show promise for cardiac regenerative therapy. PMID:25336194

Propulsion Estimates for High Energy Lunar Missions Using Future Propellants

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.; Bennett, Gary L.

2016-01-01

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 s. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

Systems engineering: A formal approach. Part 1: System concepts

NASA Astrophysics Data System (ADS)

Vanhee, K. M.

1993-03-01

Engineering is the scientific discipline focused on the creation of new artifacts that are supposed to be of some use to our society. Different types of artifacts require different engineering approaches. However, in all these disciplines the development of a new artifact is divided into stages. Three stages can always be recognized: Analysis, Design, and Realization. The book considers only the first two stages of the development process. It focuses on a specific type of artifacts, called discrete dynamic systems. These systems consist of active components of actors that consume and produce passive components or tokens. Three subtypes are studied in more detail: business systems (like a factory or restaurant), information systems (whether automated or not), and automated systems (systems that are controlled by an automated information system). The first subtype is studied by industrial engineers, the last by software engineers and electrical engineers, whereas the second is a battlefield for all three disciplines. The union of these disciplines is called systems engineering.

Pegasus delivers SLS engine section

NASA Image and Video Library

2017-03-03

NASA engineers install test hardware for the agency's new heavy lift rocket, the Space Launch System, into a newly constructed 50-foot structural test stand at NASA's Marshall Space Flight Center. In the stand, hydraulic cylinders will be electronically controlled to push, pull, twist and bend the test article with millions of pounds of force. Engineers will record and analyze over 3,000 channels of data for each test case to verify the capabilities of the engine section and validate that the design and analysis models accurately predict the amount of loads the core stage can withstand during launch and ascent. The engine section, recently delivered via NASA's barge Pegasus from NASA's Michoud Assembly Facility, is the first of four core stage structural test articles scheduled to be delivered to Marshall for testing. The engine section, located at the bottom of SLS's massive core stage, will house the rocket's four RS-25 engines and be an attachment point for the two solid rocket boosters.

Pegasus delivers SLS engine section

NASA Image and Video Library

2017-05-18

NASA engineers install test hardware for the agency's new heavy lift rocket, the Space Launch System, into a newly constructed 50-foot structural test stand at NASA's Marshall Space Flight Center. In the stand, hydraulic cylinders will be electronically controlled to push, pull, twist and bend the test article with millions of pounds of force. Engineers will record and analyze over 3,000 channels of data for each test case to verify the capabilities of the engine section and validate that the design and analysis models accurately predict the amount of loads the core stage can withstand during launch and ascent. The engine section, recently delivered via NASA's barge Pegasus from NASA's Michoud Assembly Facility, is the first of four core stage structural test articles scheduled to be delivered to Marshall for testing. The engine section, located at the bottom of SLS's massive core stage, will house the rocket's four RS-25 engines and be an attachment point for the two solid rocket boosters.

NASA on a Strong Roll in Preparing Space Launch System Flight Engines

NASA Image and Video Library

2017-08-09

NASA is on a roll when it comes to testing engines for its new Space Launch System (SLS) rocket that will send astronauts to deep-space destinations, including Mars. Just two weeks after the third test of a new RS-25 engine flight controller, the space agency recorded its fourth full-duration controller test Aug. 9 at Stennis Space Center near Bay St. Louis, Mississippi. Engineers conducted a 500-second test of the RS-25 engine controller on the A-1 Test Stand at Stennis. The test involved installing the controller on an RS-25 development engine and firing it in the same manner, and for the same length of time, as needed during an actual SLS launch. The test marked another milestone toward launch of the first integrated flight of the SLS rocket and Orion crew vehicle. Exploration Mission-1 will be an uncrewed mission into lunar orbit, designed to provide a final check-out test of rocket and Orion capabilities before astronauts are returned to deep space. The SLS rocket will be powered at launch by four RS-25 engines, providing a combined 2 million pounds of thrust, and with a pair of solid rocket boosters, providing more than 8 million pounds of total thrust. The RS-25 engines for the initial SLS flights are former space shuttle main engines that are now being used to launch the larger and heavier SLS rocket and with the new controller. The controller is a critical component that operates as the engine “brain” that communicates with SLS flight computers to receive operation performance commands and to provide diagnostic data on engine health and status. Engineers conducted early prototype tests at Stennis to collect data for development of the new controller by NASA, RS-25 prime contractor Aerojet Rocketdyne and subcontractor Honeywell. Testing of actual flight controllers began at Stennis in March. NASA is testing all controllers and engines designated for the EM-1 flight at Stennis. It also will test the SLS core stage for the flight at Stennis, which will involve installing the stage on the B-2 Test Stand and firing its four RS-25 engines simultaneously, as during an actual launch. RS-25 tests at Stennis are conducted by a team of NASA, Aerojet Rocketdyne and Syncom Space Services engineers and operators. Aerojet Rocketdyne is the RS-25 prime contractor. Syncom Space Services is the prime contractor for Stennis facilities and operations.

Energy efficient engine high pressure turbine test hardware detailed design report

NASA Technical Reports Server (NTRS)

Halila, E. E.; Lenahan, D. T.; Thomas, T. T.

1982-01-01

The high pressure turbine configuration for the Energy Efficient Engine is built around a two-stage design system. Moderate aerodynamic loading for both stages is used to achieve the high level of turbine efficiency. Flowpath components are designed for 18,000 hours of life, while the static and rotating structures are designed for 36,000 hours of engine operation. Both stages of turbine blades and vanes are air-cooled incorporating advanced state of the art in cooling technology. Direct solidification (DS) alloys are used for blades and one stage of vanes, and an oxide dispersion system (ODS) alloy is used for the Stage 1 nozzle airfoils. Ceramic shrouds are used as the material composition for the Stage 1 shroud. An active clearance control (ACC) system is used to control the blade tip to shroud clearances for both stages. Fan air is used to impinge on the shroud casing support rings, thereby controlling the growth rate of the shroud. This procedure allows close clearance control while minimizing blade tip to shroud rubs.

Vice President Pence Visits SLS Engineering Test Facility

NASA Image and Video Library

2017-09-25

The Vice President toured the SLS engineering facility where the engine section of the rocket’s massive core stage is undergoing a major stress test. The rocket’s four RS-25 engines and the two solid rocket boosters that attach to the SLS engine section will produce more than 8 million pounds of thrust to launch the Orion spacecraft beyond low-Earth orbit. More than 3,000 measurements using sensors installed on the test section will help ensure the core stage for all SLS missions can withstand the extreme forces of flight.

SRB thermal protection systems materials test results in an arc-heated nitrogen environment

NASA Technical Reports Server (NTRS)

Wojciechowski, C. J.

1979-01-01

The external surface of the Solid Rocket Booster (SRB) will experience imposed thermal and shear environments due to aerodynamic heating and radiation heating during launch, staging and reentry. This report is concerned with the performance of the various TPS materials during the staging maneuver. During staging, the wash from the Space Shuttle Main Engine (SSME) exhust plumes impose severe, short duration, thermal environments on the SRB. Five different SRB TPS materials were tested in the 1 MW Arc Plasma Generator (APG) facility. The maximum simulated heating rate obtained in the APG facility was 248 Btu/sq ft./sec, however, the test duration was such that the total heat was more than simulated. Similarly, some local high shear stress levels of 0.04 psia were not simulated. Most of the SSME plume impingement area on the SRB experiences shear stress levels of 0.02 psia and lower. The shear stress levels on the test specimens were between 0.021 and 0.008 psia. The SSME plume stagnation conditions were also simulated.

Study to Improve Airframe Turbine Engine Rotor Blade Containment

DTIC Science & Technology

1977-07-01

REPORT NO. FAA-RD-77-44 ( DOT-FA76WA-3843 JUNE 1976 STUDY TO IMPROVE AIRFRAME TURBINE ENGINE ROTOR BLADE CONTAINMENT C. 0. GUNDERSON SOF Tftj. -" So...both engines appeared to be able to marginally contain the 1 and 2 blade fragments in all compressor and turbine stages, but probably would rfiot have...adjacent blades including serrations from any stage. The investigation was made on high bypass ratio turbofan engines which power wide body transports

Animation: What makes up the Space Launch System’s massive core stage

NASA Image and Video Library

2017-04-24

NASA’s new rocket, the Space Launch System, will be the most powerful rocket ever built for deep-space missions. The 212-foot core stage is the largest rocket stage ever built and will fuel four RS-25 engines that will help launch SLS. This animation depicts the parts that make up the core stage and how these parts will be joined to form the entire stage. The five major parts include: the engine section, the hydrogen tank, the intertank, the liquid oxygen tank and the forward skirt.

Developing Evaluation Measures for the Second Stage Next Generation Engine on Evolved Expendable Launch Vehicles

DTIC Science & Technology

2012-03-01

xii THIS PAGE INTENTIONALLY LEFT BLANK xiii LIST OF ACRONYMS AND ABBREVIATIONS CDR Critical Design Review DCSS Delta Cryogenic Second Stage...seen below in Figure 5, include the Common Booster Core powered by a Pratt and Whitney Rocketdyne RS-68 engine, a Delta Cryogenic Second Stage (DCSS...do have one significant similarity. The Centaur of the Atlas V and Delta IV Cryogenic Second Stage, both use variants of the Pratt and Whitney

Maturation of enabling technologies for the next generation reignitable cryogenic upper stage

NASA Astrophysics Data System (ADS)

Mueller, Mark

Following the ESA decision in November 2008, a pre-development phase (Phase 1) of a future evolution of the Ariane 5 launcher (named Ariane 5 Midlife Evolution, A5ME) was started under Astrium Prime leadership. This upgraded version of the Ariane 5 launcher is based on an enhanced performance Upper Stage including the cryogenic re-ignitable VINCI engine. Thanks to this reignition capability, this new Upper Stage shall be "versatile" in the sense that it shall fulfil customer needs on a broader spectrum of orbits than the "standard" orbits (i.e. Geosynchronous Transfer Orbits, GTO) typically used for commercial telecommunications satellites. In order to meet the challenges of versatility, new technologies are currently being investigated. These technologies are mainly related -but not limited-to propellant management during the extended coasting phases with the related heat transfer into the tanks and the required multiple engine re-ignitions. Within the frame of the ESA Future Launchers Preparatory Programme (Period 2 Slice 1), the Cryogenic Upper Stage Technology project (CUST) aims to mature critical technologies to such a Technology Readiness Level (TRL) that they can be integrated into the baseline A5ME Upper Stage development schedule. In addition to A5ME application, these technologies can also be used on the future next generation European launcher. This paper shows the down-selection process implemented to identify the most crucial enabling technologies for a future versatile Upper Stage and gives a description of each technology finally selected for maturation in the frame of CUST. These include -amongst others-a Sandwich Common Bulkhead for the propellant tank, an external thermal insulation kit and various propellant management devices for the coasting phase. The paper also gives an overview on the related development and maturation plan including the tests to be conducted, as well as first results of the maturation activities themselves.

Educational Encounters of the Third Kind.

PubMed

Génova, Gonzalo; González, M Rosario

2017-12-01

An engineer who becomes an educator in a school of software engineering has the mission to teach how to design and construct software systems, therein applying his or her knowledge and expertise. However, due to their engineering background, engineers may forget that educating a person is not the same as designing a machine, since a machine has a well-defined goal, whilst a person is capable to self-propose his or her own objectives. The ethical implications are clear: educating a free person must leave space for creativity and self-determination in his or her own discovery of the way towards personal and professional fulfillment, which cannot consist only in achieving goals selected by others. We present here an argument that is applicable to most fields of engineering. However, the dis-analogy between educating students and programming robots may have a particular appeal to software engineers and computer scientists. We think the consideration of three different stages in the educational process may be useful to engineers when they act as educators. We claim that the three stages (instructing, training and mentoring) are essential to engineering education. In particular, education is incomplete if the third stage is not reached. Moreover, mentoring (the third stage aimed at developing creativity and self-determination) is incompatible with an educational assessment framework that considers the goals of the engineer are always given by others. In our view, then, an integral education is not only beyond programming the behavior of students, but also beyond having them reach those given goals.

J-2X engine test

NASA Image and Video Library

2011-12-01

NASA conducted a key stability test firing of the J-2X rocket engine on the A-2 Test Stand at Stennis Space Center on Dec. 1, marking another step forward in development of the upper-stage engine that will carry humans deeper into space than ever before. The J-2X will provide upper-stage power for NASA's new Space Launch System.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

The powerful J-2 engine is prominent in this photograph of a Saturn V Third Stage (S-IVB) resting on a transporter in the Manufacturing Facility at Marshall Space Flight Center in Huntsville, Alabama. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Orbit transfer vehicle engine study, phase A extension. Volume 2A: Study results

NASA Technical Reports Server (NTRS)

1980-01-01

Engine trade studies and systems analyses leading to a baseline engine selection for advanced expander cycle engine are discussed with emphasis on: (1) performance optimization of advanced expander cycle engines in the 10 to 20K pound thrust range; (2) selection of a recommended advanced expander engine configuration based on maximized performance and minimized mission risk, and definition of the components for this configuration; (3) characterization of the low thrust adaptation requirements and performance for the staged combustion engine; (4) generation of a suggested safety and reliability approach for OTV engines independent of engine cycle; (5) definition of program risk relationships between expander and staged combustion cycle engines; and (6) development of schedules and costs for the DDT&E, production, and operation phases of the 10K pound thrust expander engine program.

JT8D-100 turbofan engine, phase 1. [noise reduction

NASA Technical Reports Server (NTRS)

1974-01-01

The JT8D turbofan engine, widely used in short and medium range transport aircraft, contributes substantially to airport community noise. The jet noise is predominant in the JT8D engine and may be reduced in a modified engine, without loss of thrust, by increasing the airflow to reduce jet velocity. A configuration study evaluated the effects of fan airflow, fan pressure ratio, and bypass ratio on noise, thrust, and fuel comsumption. The cycle selected for the modified engine was based upon an increased diameter, single-stage fan and two additional core engine compressor stages, which replace the existing two-stage fan. Modifications were also made to the low pressure turbine to provide the increased torque required by the larger diameter fan. The resultant JT8D-100 engine models have the following characteristics at take-off thrust, compared to the current JT8D engine: Airflow and bypass ratio are increased, and fan pressure ratio and engine speed are reduced. The resultant engine is also longer, larger in diameter, and heavier than the JT8D base model, but these latter changes are compensated by the increased thrust and decreased fuel comsumption of the modified engine, thus providing the capability for maintaining the performance of the current JT8D-powered aircraft.

77 FR 30371 - Airworthiness Directives; International Aero Engines AG Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-23

... (USIs) of certain high-pressure compressor (HPC) stage 3 to 8 drums, and replacement of drum attachment... Condition This AD results from reports of 50 additional high-pressure compressor (HPC) stage 3 to 8 drums...

Development of the scientific heritage of M.E. Deich in the sphere of the gas dynamics of two-phase media (On the 100th anniversary of his birthday)

NASA Astrophysics Data System (ADS)

Avetisyan, A. R.; Lazarev, L. Ya.

2017-07-01

This article is a brief overview of some scientific and engineering ideas in the sphere of two-phase gas dynamics that were developed by the team of the Problem Laboratory of Turbomachines, Department of Steam and Gas Turbines, Moscow Power Engineering Institute (NRU MPEI, National Research University), under the leadership of Mikhail Efimovich Deich since 1963 and the analysis of their development and influence on the current state of the problem. At the early stages of the studies on two-phase media, the problem of the measurement of physical parameters of phases was especially urgent. The characteristics of probes for the measurement of one-phase flows in the presence of drops were studied, and the corrections for the influence of the second phase were obtained. However, the main focus was the development of new methods, and the optical method using a laser light source that is currently used at the leading laboratories of the world was chosen as the main method. The study of the wet-steam flow in nozzles is one of the first stages of the research on the problem. In these studies, the wave structure of supersonic wet-steam flows (condensation jumps and shock waves, Mach waves, turbulent condensation, periodic condensation nonstationarity, etc.) was investigated in detail. At present, like in the earlier studies, much attention is paid to the study of the influence of the addition of surface-active substance (SASs) on the wet-steam flow. The study of the wet-steam motion in steam-turbine stages was performed simultaneously with physical studies as the practical application of the obtained results. The development of computer technology in the 21st century contributed to the elaboration of the theoretical methods for the calculation of wet-steam flows in elements of power devices.

Labour perspectives of engineering degrees in the European Higher Education Area (EHEA): a case of study in the University of Cordoba (Spain)

NASA Astrophysics Data System (ADS)

Taguas, E. V.; Redel, M. D.; Pérez, R.; Peña, A.

2009-04-01

The Bologna process is reaching its final stages and is causing controversy among students. The adaptation of European universities to the European Higher Education Area (EHEA) entails not only the modification of curricular programmes and the nomenclature and duration of degrees, but also the incorporation of new teaching strategies aimed at ensuring that students acquire transversal skills and aptitudes and at increasing student participation in the teaching-learning process. A number of surveys have been carried out during the last few courses among students doing degrees in engineering (Industrial Engineering, Agronomy Engineering and Forestry Engineering). These surveys include questions on their knowledge of Bologna process, its advantages and drawbacks, their opinion about optional masters or doctorate degrees, what perspectives their degrees have on the labour market and suggestions for improvement. Although the different degrees showed notable differences, the content of EHEA is well-known by less than 30% of students, while 40% of them state they know about their perspectives on the labour market. The main advantages of EHEA were related to the improvement of practical knowledge in the subjects, the recognition of degrees in Europe and wider working opportunities. The main drawbacks pointed out were worse and shorter training periods, higher costs and fiercer competition between different degrees. In addition, they suggested that the new degrees are better adjusted to the demands of the labour market. 60% and 40% of them, respectively, approved of Masters degrees and PhDs. These features should be taken into account to organize and improve the contents of the degrees as well as to involve the students in the future of University education.

KSC-05PD-1590

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Media gather in the television studio at the NASA News Center to hear members of the Mission Management Team reveal aspects of the troubleshooting and testing being done on the liquid hydrogen tank low-level fuel cut-off sensor. On the stage at right are (from left) Bruce Buckingham, NASA news chief; Wayne Hale, Space Shuttle deputy program manager; John Muratore, manager of Systems Engineering and Integration for the Space Shuttle Program; and Mike Wetmore, director of Space Shuttle Processing. The sensor failed a routine prelaunch check during the launch countdown July 13, causing mission managers to scrub Discovery's first launch attempt. The sensor protects the Shuttle's main engines by triggering their shutdown in the event fuel runs unexpectedly low. The sensor is one of four inside the liquid hydrogen section of the External Tank (ET).

Cyclic structural analyses of anisotropic turbine blades for reusable space propulsion systems. [ssme fuel turbopump

NASA Technical Reports Server (NTRS)

Manderscheid, J. M.; Kaufman, A.

1985-01-01

Turbine blades for reusable space propulsion systems are subject to severe thermomechanical loading cycles that result in large inelastic strains and very short lives. These components require the use of anisotropic high-temperature alloys to meet the safety and durability requirements of such systems. To assess the effects on blade life of material anisotropy, cyclic structural analyses are being performed for the first stage high-pressure fuel turbopump blade of the space shuttle main engine. The blade alloy is directionally solidified MAR-M 246 alloy. The analyses are based on a typical test stand engine cycle. Stress-strain histories at the airfoil critical location are computed using the MARC nonlinear finite-element computer code. The MARC solutions are compared to cyclic response predictions from a simplified structural analysis procedure developed at the NASA Lewis Research Center.

Field Evaluation of Six Protective Coatings Applied to T56 Turbines after 1500 Hours Engine Use

DTIC Science & Technology

1991-06-01

Six Coating Systems On First-stage Gas Turbine Blades In The Engines of a Long-Range Maritime Patrol Aircraft ", Surface and Coating Technology, 36...based coatings. They were applied to the first-stage turbine blades in the engines of two long range maritime patrol aircraft operated by the Royal...incorrect. These differently coated turbine - blades have in fact seen 1500 hours service in a T56 engine . The title and further reference in the text should

J-2X Upper Stage Engine: Hardware and Testing 2009

NASA Technical Reports Server (NTRS)

Buzzell, James C.

2009-01-01

Mission: Common upper stage engine for Ares I and Ares V. Challenge: Use proven technology from Saturn X-33, RS-68 to develop the highest Isp GG cycle engine in history for 2 missions in record time . Key Features: LOX/LH2 GG cycle, series turbines (2), HIP-bonded MCC, pneumatic ball-sector valves, on-board engine controller, tube-wall regen nozzle/large passively-cooled nozzle extension, TEG boost/cooling . Development Philosophy: proven hardware, aggressive schedule, early risk reduction, requirements-driven.

Saturn Apollo Program

NASA Image and Video Library

1965-08-01

Two workers are dwarfed by the five J-2 engines of the Saturn V second stage (S-II) as they make final inspections prior to a static test firing by North American Space Division. These five hydrogen -fueled engines produced one million pounds of thrust, and placed the Apollo spacecraft into earth orbit before departing for the moon. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Reusable Agena study. Volume 1: Executive summary. [space shuttle Agena upper stage tug concept

NASA Technical Reports Server (NTRS)

1974-01-01

The shuttle Agena upper stage interim tug concept is based on a building block approach. These building block concepts are extensions of existing ascent Agena configurations. Several current improvements, have been used in developing the shuttle/Agena upper stage concepts. High-density acid is used as the Agena upper stage oxidizer. The baffled injector is used in the main engine. The DF-224 is a fourth generation computer currently in development and will be flight proven in the near future. The Agena upper stage building block concept uses the current Agena as a baseline, adds an 8.5-inch (21.6 cm) extension to the fuel tank for optimum mixture ratio, uses monomethyl hydrazine as fuel, exchanges a 150:1 nozzle extension for the existing 45:1, exchanges an Autonetics DF-224 for the existing Honeywell computer, and adds a star sensor for guidance update. These modifications to the current Agena provide a 5-foot (1.52m) diameter shuttle/Agena upper stage that will fly all Vandenberg Air Force Base missions in the reusable mode without resorting to a kick motor. The delta V velocity of the Agena is increased by use of a strap-on propellant tank option. This option provides a shuttle/Agena upper stage with the capability to place almost 3900 pounds (1769 kg) into geosynchronous orbit (24 hour period) without the aid of kick motors.

The cost of performance - A comparison of the space transportation main engine and the Space Shuttle main engine

NASA Technical Reports Server (NTRS)

Barisa, B. B.; Flinchbaugh, G. D.; Zachary, A. T.

1989-01-01

This paper compares the cost of the Space Shuttle Main Engine (SSME) and the Space Transportation Main Engine (STME) proposed by the Advanced Launch System Program. A brief description of the SSME and STME engines is presented, followed by a comparison of these engines that illustrates the impact of focusing on acceptable performance at minimum cost (as for the STME) or on maximum performance (as for the SSME). Several examples of cost reduction methods are presented.

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine

NASA Technical Reports Server (NTRS)

Kopasakis, George; Connolly, Joseph W.; Cheng, Larry

2015-01-01

This paper covers the development of stage-by-stage and parallel flow path compressor modeling approaches for a Variable Cycle Engine. The stage-by-stage compressor modeling approach is an extension of a technique for lumped volume dynamics and performance characteristic modeling. It was developed to improve the accuracy of axial compressor dynamics over lumped volume dynamics modeling. The stage-by-stage compressor model presented here is formulated into a parallel flow path model that includes both axial and rotational dynamics. This is done to enable the study of compressor and propulsion system dynamic performance under flow distortion conditions. The approaches utilized here are generic and should be applicable for the modeling of any axial flow compressor design.

Near Stall Flow Analysis in the Transonic Fan of the RTA Propulsion System

NASA Technical Reports Server (NTRS)

Hah, Chunill

2010-01-01

Turbine-based propulsion systems for access to space have been investigated at NASA Glenn Research center. A ground demonstrator engine for validation testing has been developed as a part of the program. The demonstrator, the Revolutionary Turbine Accelerator (RTA-1), is a variable cycle turbofan ramjet designed to transition from an augmented turbofan to a ramjet that produces the thrust required to accelerate the vehicle to Mach 4. The RTA-1 is designed to accommodate a large variation in bypass ratio from sea level static to Mach 4 flight condition. A key component of this engine is a new fan stage that accommodates these large variations in bypass ratio and flow ranges. In the present study, unsteady flow behavior in the fan of the RTA-1 is studied in detail with large eddy simulation (LES) and the numerical results are compared with measured data. During the experimental study of the fan stage, humming sound was detected at 100 % speed near stall operation. The main purpose of the study is to investigate details of the unsteady flow behavior at near stall operation and to identify a possible cause of the hum. The large eddy simulation of the current flow field reproduces main features of the measured flow very well. The LES simulation indicates that non-synchronous flow instability develops as the fan operates toward the stall limit. The FFT analysis of the calculated wall pressure shows that the rotating flow instability has the characteristic frequency that is about 50% of the blade passing frequency.

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.

Propulsion Progress for NASA's Space Launch System

NASA Technical Reports Server (NTRS)

May, Todd A.; Lyles, Garry M.; Priskos, Alex S.; Kynard, Michael H.; Lavoie, Anthony R.

2012-01-01

Leaders from NASA's Space Launch System (SLS) will participate in a panel discussing the progress made on the program's propulsion systems. The SLS will be the nation's next human-rated heavy-lift vehicle for new missions beyond Earth's orbit. With a first launch slated for 2017, the SLS Program is turning plans into progress, with the initial rocket being built in the U.S.A. today, engaging the aerospace workforce and infrastructure. Starting with an overview of the SLS mission and programmatic status, the discussion will then delve into progress on each of the primary SLS propulsion elements, including the boosters, core stage engines, upper stage engines, and stage hardware. Included will be a discussion of the 5-segment solid rocket motors (ATK), which are derived from Space Shuttle and Ares developments, as well as the RS-25 core stage engines from the Space Shuttle inventory and the J- 2X upper stage engine now in testing (Pratt and Whitney Rocketdyne). The panel will respond to audience questions about this important national capability for human and scientific space exploration missions.

Heat Transfer Due to Unsteady Effects as Investigated in a High-Speed, Full-Scale, Fully-Cooled Turbine Vane and Rotor Stage

DTIC Science & Technology

2008-06-01

the turbine stages of these engines have been established. The predominant method of cooling vane and rotor airfoils , having been used for over...INVESTIGATED IN A HIGH-SPEED, FULL-SCALE, FULLY-COOLED TURBINE VANE AND ROTOR STAGE THESIS Presented to the Faculty Department of Aeronautics and...reduce the effectiveness of film cooling in the vane and rotor stages of turbine engines . Even today, fairly little experimentation has been

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

This photograph shows how the fuel tank assembly and the liquid oxygen tank for the Saturn V S-IC (first) stage are placed side by side prior to commencement of the mating of the two stages in the Marshall Space Flight Center, building 4705. The fuel tank carried kerosene as its fuel. The S-IC stage used five F-1 engines, that used kerosene and liquid oxygen as propellant and each engine provided 1,500,000 pounds of thrust. This stage lifted the entire vehicle and Apollo spacecraft from the launch pad.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A J-2 engine undergoes static firing. The J-2, developed under the direction of the Marshall Space Flight Center, was propelled by liquid hydrogen and liquid oxygen. A single J-2 was utilized in the S-IVB stage (the second stage for the Saturn IB and third stage for the Saturn V) and in a cluster of five for the second stage (S-II) of the Saturn V. Initially rated at 200,000 pounds of thrust, the engine was later uprated in the Saturn V program to 230,000 pounds.

Design and Testing of the Contra-Rotating Turbine for the Scimitar Precooled Mach 5 Cruise Engine

NASA Astrophysics Data System (ADS)

Varvill, R.; Paniagua, G.; Kato, H.; Thatcher, M.

tion chamber and subsequent expansion through the main noz- zle to produce thrust. In subsonic flight it becomes the gas generator driving a high bypass ratio ducted fan through a hub turbine, the exhaust mixing with the duct flow and discharging through the bypass nozzle to produce thrust. In both modes the turbo-compressor is driven by a helium turbine which has contra rotating stages to improve its efficiency at low rotational speed and reduce the number of stages required. Due to the large speed of sound mismatch between the air compressor and the helium turbine it is possible to eliminate the turbine stators by contra rotating the spools. The compressor is divided into low pressure and high pressure spools although by normal gas turbine standards they are both low pressure ratio machines.

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.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This small group of unidentified officials is dwarfed by the gigantic size of the Saturn V first stage (S-1C) at the shipping area of the Manufacturing Engineering Laboratory at Marshall Space Flight Center in Huntsville, Alabama. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

This is a cutaway view of the Saturn V first stage, known as the S-IC, detailing the five F-1 engines and fuel cells. The S-IC stage is 138 feet long and 33 feet in diameter, producing more than 7,500,000 pounds of thrust through the five F-1 engines that are powered by liquid oxygen and kerosene. Four of the engines are mounted on an outer ring and gimbal for control purposes. The fifth engine is rigidly mounted in the center. When ignited, the roar produced by the five engines equals the sound of 8,000,000 hi-fi sets.

Evaluation of a Game to Teach Requirements Collection and Analysis in Software Engineering at Tertiary Education Level

ERIC Educational Resources Information Center

Hainey, Thomas; Connolly, Thomas M.; Stansfield, Mark; Boyle, Elizabeth A.

2011-01-01

A highly important part of software engineering education is requirements collection and analysis which is one of the initial stages of the Database Application Lifecycle and arguably the most important stage of the Software Development Lifecycle. No other conceptual work is as difficult to rectify at a later stage or as damaging to the overall…

77 FR 60288 - Airworthiness Directives; Pratt & Whitney (P&W) Division Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-10-03

... reassembly of the 2nd stage high-pressure turbine (HPT) rotor and 3rd stage LPT rotor at the next HPT and LPT... found. (2) At the next high-pressure turbine (HPT) overhaul, re- assemble the 2nd stage HPT rotor blades... fractures in the low-pressure turbine (LPT) of certain PW4000-94'' and PW4000-100'' turbofan engines. This...

Saturn V First Stage Leaves the Dynamic Test Stand

NASA Technical Reports Server (NTRS)

1967-01-01

This photo shows the Saturn V first stage being lowered to the ground following a successful test to determine the effects of continual vibrations simulating the effects of an actual launch. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

77 FR 40479 - Airworthiness Directives; Rolls-Royce Corporation Turboshaft Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-07-10

... inspection (FPI) on certain 3rd and 4th stage turbine wheels for cracks in the turbine blades. We are issuing this AD to prevent failure of 3rd or 4th stage turbine wheel blades which could cause engine failure... certain 3rd and 4th stage turbine wheels for cracks in the turbine blades. Comments We gave the public the...

77 FR 4648 - Airworthiness Directives; Rolls-Royce plc (RR) RB211-535 Series Turbofan Engine

Federal Register 2010, 2011, 2012, 2013, 2014

2012-01-31

... inspections (FPI) of the low-pressure (LP) turbine stage 1, 2, and 3 discs to detect cracks in the discs. This... turbine stage 1, 2, and 3 discs, which could result in an uncontained release of LP turbine blades and... require performing an initial FPI on the LP turbine stage 1, 2, and 3 discs at the next engine shop...

The AC-120: The advanced commercial transport

NASA Technical Reports Server (NTRS)

Duran, David; Griffin, Ernest; Mendoza, Saul; Nguyen, Son; Pickett, Tim; Noernberg, Clemm

1993-01-01

The main objective of this design was to fulfill a need for a new airplane to replace the aging 100 to 150 passenger, 1500 nautical mile range aircraft such as the Douglas DC9 and Boeing 737-100 airplanes. After researching the future aircraft market, conducting extensive trade studies, and analysis on different configurations, the AC-120 Advanced Commercial Transport final design was achieved. The AC-120's main design features include the incorporation of a three lifting surface configuration which is powered by two turboprop engines. The AC-120 is an economically sensitive aircraft which meets the new FM Stage Three noise requirements, and has lower NO(x) emissions than current turbofan powered airplanes. The AC-120 also improves on its contemporaries in passenger comfort, manufacturing, and operating cost.

Experience and prospects of oil shale utilization for power production in Russia

NASA Astrophysics Data System (ADS)

Potapov, O. P.

2016-09-01

Due to termination of work at the Leningrad Shale Deposit, the Russian shale industry has been liquidated, including not only shale mining and processing but also research and engineering (including design) activities, because this deposit was the only commercially operated complex in Russia. UTT-3000 plants with solid heat carrier, created mainly by the Russian specialists under scientific guidance of members of Krzhizhanovsky Power Engineering Institute, passed under the control of Estonian engineers, who, alongside with their operation in Narva, construct similar plants in Kohtla-Jarve, having renamed the Galoter Process into the Enifit or Petroter. The main idea of this article is to substantiate the expediency of revival of the oil shale industry in Russia. Data on the UTT-3000 plants' advantages, shale oils, and gas properties is provided. Information on investments in an UTT-3000 plant and estimated cost of Leningrad oil shale mining at the Mezhdurechensk Strip Mine is given. For more detailed technical and economic assessment of construction of a complex for oil shale extraction and processing, it is necessary to develop a feasibility study, which should be the first stage of this work. Creation of such a complex will make it possible to produce liquid and gaseous power fuel from oil shale of Leningrad Deposit and provide the opportunity to direct for export the released volumes of oil and gas for the purposes of Russian budget currency replenishment.

Application of dual-fuel propulsion to a single stage AMLS vehicle

NASA Technical Reports Server (NTRS)

Lepsch, Roger A., Jr.; Stanley, Douglas O.; Unal, Resit

1993-01-01

As part of NASA's Advanced Manned Launch System (AMLS) study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined. Several dual-fuel propulsion concepts were investigated. These include: a separate engine concept combining Russian RD-170 kerosene-fueled engines with SSME-derivative engines; the kerosene and hydrogen-fueled Russian RD-701 engine concept; and a dual-fuel, dual-expander engine concept. Analysis to determine vehicle weight and size characteristics was performed using conceptual level design techniques. A response surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicle concepts with respect to several important propulsion system and vehicle design parameters in order to achieve minimum empty weight. Comparisons were then made with a hydrogen-fueled reference, single-stage vehicle. The tools and methods employed in the analysis process are also summarized.

Ceramic high pressure gas path seal

NASA Technical Reports Server (NTRS)

Liotta, G. C.

1987-01-01

Stage 1 ceramic shrouds (high pressure turbine gas path seal) were developed for the GE T700 turbine helicopter engine under the Army/NASA Contract NAS3-23174. This contract successfully proved the viability and benefits of a Stage 1 ceramic shroud for production application. Stage 1 ceramic shrouds were proven by extensive component and engine testing. This Stage 1 ceramic shroud, plasma sprayed ceramic (ZrOs-BY2O3) and bond coating (NiCrAlY) onto a cast metal backing, offers significant engine performance improvement. Due to the ceramic coating, the amount of cooling air required is reduced 20% resulting in a 0.5% increase in horsepower and a 0.3% decrease in specific fuel consumption. This is accomplished with a component which is lower in cost than the current production shroud. Stage 1 ceramic shrouds will be introduced into field service in late 1987.

Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio. Part 2; Comparison of CFD and Experimental Results

NASA Technical Reports Server (NTRS)

Celestina, Mark L.; Suder, Kenneth L.; Kulkarni, Sameer

2010-01-01

NASA and GE teamed to design and build a 57 percent engine scaled fan stage for a Mach 4 variable cycle turbofan/ramjet engine for access to space with multipoint operations. This fan stage was tested in NASA's transonic compressor facility. The objectives of this test were to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off; 2) transition over large swings in fan bypass ratio; 3) transition from turbofan to ramjet; and 4) fan wind-milling operation at high Mach flight conditions. This paper will focus on an assessment of APNASA, a multistage turbomachinery analysis code developed by NASA, to predict the fan stage performance and operability over a wide range of speeds (37 to 100 percent) and bypass ratios.

Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

NASA Technical Reports Server (NTRS)

Shi, John J.

2005-01-01

During the development stage, in order to design/to size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads must be defined. There are two kinds of dynamic environment, i.e. shock transients and steady-state random and sinusoidal vibration environments. Usually, the steady-state random and sinusoidal vibration environments are scalable, but the shock environments are not scalable. In other words, based on similarities only random vibration environments can be defined for a new engine. The methodology covered in this paper provides a way to predict the shock environments and the dynamic loads for new engine systems and new engine components in the early stage of new engine development or engine nozzle modifications.

Fan Noise Test Facility

NASA Image and Video Library

1969-01-21

The Fan Noise Test Facility built at the Lewis Research Center to obtain far-field noise data for the National Aeronautics and Space Administration (NASA) and General Electric Quiet Engine Program. The engine incorporated existing noise reduction methods into an engine of similar power to those that propelled the Boeing 707 or McDonnell-Douglas DC-8 airliner. The new the low-bypass ratio turbofan engines of the 1960s were inherently quieter than their turbojet counterparts, researchers had a better grasp of the noise generation problem, and new acoustic technologies had emerged. Lewis contracted General Electric in 1969 to build and aerodynamically test three experimental engines with 72-inch diameter fans. The engines were then brought to Lewis and tested with an acoustically treated nacelle. This Fan Noise Test Facility was built off of the 10- by 10-Foot Supersonic Wind Tunnel’s Main Compressor and Drive Building. Lewis researchers were able to isolate the fan’s noise during these initial tests by removing the core of the engine. The Lewis test rig drove engines to takeoff tip speeds of 1160 feet per second. The facility was later used to test a series of full-scale model fans and fan noise suppressors to be used with the quiet engine. NASA researchers predicted low-speed single-stage fans without inlet guide vanes and with large spacing between rotors and stators would be quieter. General Electric modified a TF39 turbofan engine by removing the the outer protion of the fan and spacing the blade rows of the inner portion. The tests revealed that the untreated version of the engine generated less noise than was anticipated, and the acoustically treated nacelle substantially reduced engine noise.

Stennis certifies final shuttle engine

NASA Image and Video Library

2008-10-22

Steam blasts out of the A-2 Test Stand at Stennis Space Center on Oct. 22 as engineers begin a certification test on engine 2061, the last space shuttle main flight engine scheduled to be built. Since 1975, Stennis has tested every space shuttle main engine used in the program - about 50 engines in all. Those engines have powered more than 120 shuttle missions - and no mission has failed as a result of engine malfunction. For the remainder of 2008 and throughout 2009, Stennis will continue testing of various space shuttle main engine components.

X-33 Attitude Control Using the XRS-2200 Linear Aerospike Engine

NASA Technical Reports Server (NTRS)

Hall, Charles E.; Panossian, Hagop V.

1999-01-01

The Vehicle Control Systems Team at Marshall Space Flight Center, Structures and Dynamics Laboratory, Guidance and Control Systems Division is designing, under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control systems for the X-33 experimental vehicle. Test flights, while suborbital, will achieve sufficient altitudes and Mach numbers to test Single Stage To Orbit, Reusable Launch Vehicle technologies. Ascent flight control phase, the focus of this paper, begins at liftoff and ends at linear aerospike main engine cutoff (MECO). The X-33 attitude control system design is confronted by a myriad of design challenges: a short design cycle, the X-33 incremental test philosophy, the concurrent design philosophy chosen for the X-33 program, and the fact that the attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems. Additionally, however, and of special interest, the use of the linear aerospike engine is a departure from the gimbaled engines traditionally used for thrust vector control (TVC) in launch vehicles and poses certain design challenges. This paper discusses the unique problem of designing the X-33 attitude control system with the linear aerospike engine, requirements development, modeling and analyses that verify the design.

The SKYLON Spaceplane

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

SKYLON is a single stage to orbit (SSTO) winged spaceplane designed to give routine low cost access to space. At a gross takeoff weight of 275 tonnes of which 220 tonnes is propellant the vehicle is capable of placing 12 tonnes into an equatorial low Earth orbit. The vehicle configuration consists of a slender fuselage containing the propellant tankage and payload bay with delta wings located midway along the fuselage carrying the SABRE engines in axisymmetric nacelles on the wingtips. The vehicle takes off and lands horizontally on it's own undercarriage. The fuselage is constructed as a multilayer structure consisting of aeroshell, insulation, structure and tankage. SKYLON employs extant or near term materials technology in order to minimise development cost and risk. The SABRE engines have a dual mode capability. In rocket mode the engine operates as a closed cycle liquid oxygen/liquid hydrogen high specific impulse rocket engine. In airbreathing mode (from takeoff to Mach 5) the liquid oxygen flow is replaced by atmospheric air, increasing the installed specific impulse 3-6 fold. The airflow is drawn into the engine via a 2 shock axisymmetric intake and cooled to cryogenic temperatures prior to compression. The hydrogen fuel flow acts as a heat sink for the closed cycle helium loop before entering the main combustion chamber.

Lube oil-dependent ash chemistry on soot oxidation reactivity in a gasoline direct-injection engine

DOE PAGES

Choi, Seungmok; Seong, Heeje

2016-09-30

Gasoline particulate filters (GPF) are considered an enabling technology to meet stringent particulate matter (PM) regulations for gasoline direct-injection (GDI) engines, which are known to produce significant PM emissions. While ash loading in filters has been recognized to be detrimental in filter performance by increasing back pressure, increased ash fractions in soot were observed to enhance soot oxidation. In this study, GDI soot samples derived from different gasoline/lube oil blends were evaluated to identify potential promoting factors when formulated lube oils were dosed into gasoline fuel. Ca-derived ash enhanced soot oxidation remarkably, while P- and ZDDP-derived ash deteriorated soot oxidation.more » It is apparent that the promoting effect of lube oil-derived ash is due mainly to the Ca component that is the most abundant among additive components in lube oil. Bulk and surface analyses of these ash compounds indicate that Ca-derived ash would be complex compounds, while the contribution of CaSO 4, which is one of the most abundant ash compounds from diesel engines, is almost negligible. For the validation of the ash promoting impact in filters, the regeneration experiments were compared for a TWC-coated GPF in a GDI engine before and after ash loading was performed. The pressure drop of the ash-loaded GPF decreased noticeably in the initial regeneration stage and it increased gradually, whereas that of no ash-loaded GPF increased gradually without any reduction. So, it is concluded that the ash layer in the GPF assisted soot oxidation in the early regeneration stage when it was in close contact with soot.« less

Ciliates learn to diagnose and correct classical error syndromes in mating strategies

PubMed Central

Clark, Kevin B.

2013-01-01

Preconjugal ciliates learn classical repetition error-correction codes to safeguard mating messages and replies from corruption by “rivals” and local ambient noise. Because individual cells behave as memory channels with Szilárd engine attributes, these coding schemes also might be used to limit, diagnose, and correct mating-signal errors due to noisy intracellular information processing. The present study, therefore, assessed whether heterotrich ciliates effect fault-tolerant signal planning and execution by modifying engine performance, and consequently entropy content of codes, during mock cell–cell communication. Socially meaningful serial vibrations emitted from an ambiguous artificial source initiated ciliate behavioral signaling performances known to advertise mating fitness with varying courtship strategies. Microbes, employing calcium-dependent Hebbian-like decision making, learned to diagnose then correct error syndromes by recursively matching Boltzmann entropies between signal planning and execution stages via “power” or “refrigeration” cycles. All eight serial contraction and reversal strategies incurred errors in entropy magnitude by the execution stage of processing. Absolute errors, however, subtended expected threshold values for single bit-flip errors in three-bit replies, indicating coding schemes protected information content throughout signal production. Ciliate preparedness for vibrations selectively and significantly affected the magnitude and valence of Szilárd engine performance during modal and non-modal strategy corrective cycles. But entropy fidelity for all replies mainly improved across learning trials as refinements in engine efficiency. Fidelity neared maximum levels for only modal signals coded in resilient three-bit repetition error-correction sequences. Together, these findings demonstrate microbes can elevate survival/reproductive success by learning to implement classical fault-tolerant information processing in social contexts. PMID:23966987

Lube oil-dependent ash chemistry on soot oxidation reactivity in a gasoline direct-injection engine

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

Choi, Seungmok; Seong, Heeje

Gasoline particulate filters (GPF) are considered an enabling technology to meet stringent particulate matter (PM) regulations for gasoline direct-injection (GDI) engines, which are known to produce significant PM emissions. While ash loading in filters has been recognized to be detrimental in filter performance by increasing back pressure, increased ash fractions in soot were observed to enhance soot oxidation. In this study, GDI soot samples derived from different gasoline/lube oil blends were evaluated to identify potential promoting factors when formulated lube oils were dosed into gasoline fuel. Ca-derived ash enhanced soot oxidation remarkably, while P- and ZDDP-derived ash deteriorated soot oxidation.more » It is apparent that the promoting effect of lube oil-derived ash is due mainly to the Ca component that is the most abundant among additive components in lube oil. Bulk and surface analyses of these ash compounds indicate that Ca-derived ash would be complex compounds, while the contribution of CaSO 4, which is one of the most abundant ash compounds from diesel engines, is almost negligible. For the validation of the ash promoting impact in filters, the regeneration experiments were compared for a TWC-coated GPF in a GDI engine before and after ash loading was performed. The pressure drop of the ash-loaded GPF decreased noticeably in the initial regeneration stage and it increased gradually, whereas that of no ash-loaded GPF increased gradually without any reduction. So, it is concluded that the ash layer in the GPF assisted soot oxidation in the early regeneration stage when it was in close contact with soot.« less

Implanted component faults and their effects on gas turbine engine performance

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

MacLeod, J.D.; Taylor, V.; Laflamme, J.C.G.

Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effect is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library, which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine,more » such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: (a) first-stage turbine nozzle erosion damage; (b) first-stage turbine rotor blade untwist; (c) compressor seal wear; (d) first and second-stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.« less

Vacuum application of thermal barrier plasma coatings

NASA Technical Reports Server (NTRS)

Holmes, R. R.; Mckechnie, T. N.

1988-01-01

Coatings are presently applied to Space Shuttle Main Engine (SSME) turbine blades for protection against the harsh environment realized in the engine during lift off-to-orbit. High performance nickel, chromium, aluminum, and yttrium (NiCrAlY) alloy coatings, which are applied by atmospheric plasma spraying, crack and spall off because of the severe thermal shock experienced during start-up and shut-down of the engine. Ceramic coatings of yttria stabilized zirconia (ZrO2-Y2O3) were applied initially as a thermal barrier over coating to the NiCrAlY but were removed because of even greater spalling. Utilizing a vacuum plasma spraying process, bond coatings of NiCrAlY were applied in a low pressure atmosphere of argon/helium, producing significantly improved coating-to-blade bonding. The improved coatings showed no spalling after 40 MSFC burner rig thermal shock cycles, cycling between 1700 and -423 F. The current atmospheric plasma NiCrAlY coatings spalled during 25 test cycles. Subsequently, a process was developed for applying a durable thermal barrier coating of ZrO2-Y2O3 to the turbine blades of first stage high-pressure fuel turbopumps utilizing the vacuum plasma process. The improved thermal barrier coating has successfully passed 40 burner rig thermal shock cycles without spalling. Hot firing in an SSME turbine engine is scheduled for the blades. Tooling was installed in preparation for vacuum plasma spray coating other SSME hardware, e.g., the titanium main fuel valve housing (MFVH) and the fuel turbopump nozzle/stator.

NASA X-34 Technology in Motion

NASA Technical Reports Server (NTRS)

Beech, Geoffrey; Chandler, Kristie

1997-01-01

The X-34 technology development program is a joint industry/government project to develop, test, and operate a small, fully-reusable hypersonic flight vehicle. The objective is to demonstrate key technologies and operating concepts applicable to future reusable launch vehicles. Integrated in the vehicle are various systems to assure successful completion of mission objectives, including the Main Propulsion System (MPS). NASA-Marshall Space Flight Center (MSFC) is responsible for developing the X-34's MPS including the design and complete build package for the propulsion system components. The X-34 will be powered by the Fastrac Engine, which is currently in design and development at NASA-MSFC. Fastrac is a single-stage main engine, which burns a mixture of liquid oxygen (LOX) and kerosene(RP-1). The interface between the MPS and Fastrac engine are critical for proper system operation and technologies applicable to future reusable launch vehicles. Deneb's IGRIP software package with the Dynamic analysis option provided a key tool for conducting studies critical to this interface as well as a mechanism to drive the design of the LOX and RP-1 feedlines. Kinematic models were created for the Fastrac Engine and the feedlines for various design concepts. Based on the kinematic simulation within Envision, design and joint limits were verified and system interference controlled. It was also critical to the program to evaluate the effect of dynamic loads visually, providing a verification tool for dynamic analysis and in some cases uncovering areas that had not been considered. Deneb's software put the X-34 technology in motion and has been a key factor in facilitating the strenuous design schedule.

High/variable mixture ratio O2/H2 engine

NASA Technical Reports Server (NTRS)

Adams, A.; Parsley, R. C.

1988-01-01

Vehicle/engine analysis studies have identified the High/Dual Mixture Ratio O2/H2 Engine cycle as a leading candidate for an advanced Single Stage to Orbit (SSTO) propulsion system. This cycle is designed to allow operation at a higher than normal O/F ratio of 12 during liftoff and then transition to a more optimum O/F ratio of 6 at altitude. While operation at high mixture ratios lowers specific impulse, the resultant high propellant bulk density and high power density combine to minimize the influence of atmospheric drag and low altitude gravitational forces. Transition to a lower mixture ratio at altitude then provides improved specific impulse relative to a single mixture ratio engine that must select a mixture ratio that is balanced for both low and high altitude operation. This combination of increased altitude specific impulse and high propellant bulk density more than offsets the compromised low altitude performance and results in an overall mission benefit. Two areas of technical concern relative to the execution of this dual mixture ratio cycle concept are addressed. First, actions required to transition from high to low mixture ratio are examined, including an assessment of the main chamber environment as the main chamber mixture ratio passes through stoichiometric. Secondly, two approaches to meet a requirement for high turbine power at high mixture ratio condition are examined. One approach uses high turbine temperature to produce the power and requires cooled turbines. The other approach incorporates an oxidizer-rich preburner to increase turbine work capability via increased turbine mass flow.

Tissue engineering, stem cells, and cloning for the regeneration of urologic organs.

PubMed

Atala, Anthony

2003-10-01

Tissue engineering efforts are currently being undertaken for every type of tissue and organ within the urinary system. Most of the effort expended to engineer genitourinary tissues has occurred within the last decade. Tissue engineering techniques require a cell culture facility designed for human application. Personnel who have mastered the techniques of cell harvest, culture, and expansion as well as polymer design are essential for the successful application of this technology. Various engineered genitourinary tissues are at different stages of development, with some already being used clinically, a few in preclinical trials, and some in the discovery stage. Recent progress suggests that engineered urologic tissues may have an expanded clinical applicability in the future.

75 FR 14379 - Airworthiness Directives; Rolls-Royce Deutschland Ltd & Co KG (RRD) Models Tay 620-15, Tay 650-15...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-25

...: Following a review of operational data of the Tay 651-54 engine, it has been found that the actual stress... found that the actual stress levels in the Tay 651-54 engine High Pressure Compressor (HPC) stages 1, 3... actual stress levels in the Tay 651-54 engine High Pressure Compressor (HPC) stages 1, 3, 6, 7 and 12...

Closed-loop air cooling system for a turbine engine

DOEpatents

North, William Edward

2000-01-01

Method and apparatus are disclosed for providing a closed-loop air cooling system for a turbine engine. The method and apparatus provide for bleeding pressurized air from a gas turbine engine compressor for use in cooling the turbine components. The compressed air is cascaded through the various stages of the turbine. At each stage a portion of the compressed air is returned to the compressor where useful work is recovered.

Liquid Oxygen/Liquid Methane Propulsion and Cryogenic Advanced Development

NASA Technical Reports Server (NTRS)

Klem, Mark D.; Smith, Timothy D.; Wadel, Mary F.; Meyer, Michael L.; Free, James M.; Cikanek, Harry A., III

2011-01-01

Exploration Systems Architecture Study conducted by NASA in 2005 identified the liquid oxygen (LOx)/liquid methane (LCH4) propellant combination as a prime candidate for the Crew Exploration Vehicle Service Module propulsion and for later use for ascent stage propulsion of the lunar lander. Both the Crew Exploration Vehicle and Lunar Lander were part the Constellation architecture, which had the objective to provide global sustained lunar human exploration capability. From late 2005 through the end of 2010, NASA and industry matured advanced development designs for many components that could be employed in relatively high thrust, high delta velocity, pressure fed propulsion systems for these two applications. The major investments were in main engines, reaction control engines, and the devices needed for cryogenic fluid management such as screens, propellant management devices, thermodynamic vents, and mass gauges. Engine and thruster developments also included advanced high reliability low mass igniters. Extensive tests were successfully conducted for all of these elements. For the thrusters and engines, testing included sea level and altitude conditions. This advanced development provides a mature technology base for future liquid oxygen/liquid methane pressure fed space propulsion systems. This paper documents the design and test efforts along with resulting hardware and test results.

Monitoring/Imaging and Regenerative Agents for Enhancing Tissue Engineering Characterization and Therapies.

PubMed

Santiesteban, Daniela Y; Kubelick, Kelsey; Dhada, Kabir S; Dumani, Diego; Suggs, Laura; Emelianov, Stanislav

2016-03-01

The past three decades have seen numerous advances in tissue engineering and regenerative medicine (TERM) therapies. However, despite the successes there is still much to be done before TERM therapies become commonplace in clinic. One of the main obstacles is the lack of knowledge regarding complex tissue engineering processes. Imaging strategies, in conjunction with exogenous contrast agents, can aid in this endeavor by assessing in vivo therapeutic progress. The ability to uncover real-time treatment progress will help shed light on the complex tissue engineering processes and lead to development of improved, adaptive treatments. More importantly, the utilized exogenous contrast agents can double as therapeutic agents. Proper use of these Monitoring/Imaging and Regenerative Agents (MIRAs) can help increase TERM therapy successes and allow for clinical translation. While other fields have exploited similar particles for combining diagnostics and therapy, MIRA research is still in its beginning stages with much of the current research being focused on imaging or therapeutic applications, separately. Advancing MIRA research will have numerous impacts on achieving clinical translations of TERM therapies. Therefore, it is our goal to highlight current MIRA progress and suggest future research that can lead to effective TERM treatments.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

The fuel tank assembly of the Saturn V S-IC (first) stage is readied to be mated to the liquid oxygen tank at the Marshall Space Flight Center. The fuel tank carried kerosene as its fuel. The S-IC stage utilized five F-1 engines that used kerosene and liquid oxygen as propellant. Each engine provided 1,500,000 pounds of thrust. This stage lifted the entire vehicle and Apollo spacecraft from the launch pad.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

The fuel tank assembly for the Saturn V S-IC (first) stage arrived at the Marshall Space Flight Center, building 4707, for mating to the liquid oxygen tank. The fuel tank carried kerosene as its fuel. The S-IC stage used five F-1 engines, that used kerosene and liquid oxygen as propellant and each engine provided 1,500,000 pounds of thrust. This stage lifted the entire vehicle and Apollo spacecraft from the launch pad.

Analysis of middle bearing failure in rotor jet engine using tip-timing and tip-clearance techniques

NASA Astrophysics Data System (ADS)

Rzadkowski, R.; Rokicki, E.; Piechowski, L.; Szczepanik, R.

2016-08-01

The reported problem is the failure of the middle bearing in an aircraft rotor engine. Tip-timing and tip-clearance and variance analyses are carried out on a compressor rotor blade in the seventh stage above the middle bearing. The experimental analyses concern both an aircraft engine with a middle bearing in good working order and an engine with a damaged middle bearing. A numerical analysis of seventh stage blade free vibration is conducted to explain the experimental results. This appears to be an effective method of predicting middle bearing failure. The results show that variance first increases in the initial stages of bearing failure, but then starts to decrease and stabilize, and then again decrease shortly before complete bearing failure.

Dr. von Braun Standing by Five F-1 Engines

NASA Technical Reports Server (NTRS)

2004-01-01

A pioneer of America's space program, Dr. von Braun stands by the five F-1 engines of the Saturn V launch vehicle. This Saturn V vehicle is an actual test vehicle which has been displayed at the U.S. Space Rocket Center in Huntsville, Alabama. Designed and developed by Rocketdyne under the direction of the Marshall Space Flight Center, a cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage. The engines measured 19-feet tall by 12.5-feet at the nozzle exit and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust. The S-IC stage is the first stage, or booster, of a 364-foot long rocket that ultimately took astronauts to the Moon.

Use of bioreactors in maxillofacial tissue engineering.

PubMed

Depprich, Rita; Handschel, Jörg; Wiesmann, Hans-Peter; Jäsche-Meyer, Janine; Meyer, Ulrich

2008-07-01

Engineering of various oral tissues is a challenging issue in contemporary maxillofacial reconstructive research. In contrast to the classic biomaterial approach, tissue engineering is based on the understanding of cell driven tissue formation, and aims to generate new functional tissues, rather than just to implant non-living space holders. Researchers hope to reach this goal by combining knowledge from biology, physics, materials science, engineering, and medicine in an integrated manner. Several major technical advances have been made in this field during the last decade, and clinical application is at the stage of first clinical trials. A recent limitation of extracorporally engineered cellular substitutes is the problem of growing enlarged tissues ex vivo. One of the main research topics is therefore to scale up artificial tissue constructs for use in extended defect situations. To overcome the monolayer inherent two-dimensional cell assembly, efforts have been made to grow cells in a three-dimensional space. Bioreactors have therefore been in focus for a considerable time to build up enlarged tissues. The shift from the ex vivo approach of cell multiplication to the generation of a real tissue growth is mirrored by the development of bioreactors, enabling scientists to grow more complex tissue constructs. This present review intends to provide an overview of the current state of art in maxillofacial tissue engineering by the use of bioreactors, its limitations and hopes, as well as the future research trends.

Upper-stage space shuttle propulsion by means of separate scramjet and rocket engines

NASA Technical Reports Server (NTRS)

Franciscus, L. C.; Allen, J. L.

1972-01-01

A preliminary mission study of a reusable vehicle from staging to orbit indicates payload advantages for a dual-propulsion system consisting of separate scramjet and rocket engines. In the analysis the scramjet operated continuously and the initiation of rocket operation was varied. For a stage weight of 500,000 lb the payload was 10.4 percent of stage weight or 70 percent greater than that of a comparable all-rocket-powered stage. When compared with a reusable two-state rocket vehicle having 50,000 lb payload, the use of the dual propulsion system for the second stage resulted in significant decreases in lift-off weight and empty weight, indicating possible lower hardware costs.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

This cutaway illustration shows the Saturn V S-IC (first) stage with detailed callouts of the components. The S-IC Stage is 138 feet long and 33 feet in diameter, producing 7,500,000 pounds of thrust through five F-1 engines that are powered by liquid oxygen and kerosene. Four of the engines are mounted on an outer ring and gimbal for control purposes. The fifth engine is rigidly mounted in the center. When ignited, the roar produced by the five engines equals the sound of 8,000,000 hi-fi sets.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This illustration shows a cutaway drawing with callouts of the major components for the S-IC (first) stage of the Saturn V launch vehicle. The S-IC stage is 138 feet long and 33 feet in diameter, producing more than 7,500,000 pounds of thrust through five F-1 engines powered by liquid oxygen and kerosene. Four of the engines are mounted on an outer ring and gimball for control purposes. The fifth engine is rigidly mounted in the center. When ignited, the roar produced by the five engines equals the sound of 8,000,000 hi-fi sets.

Regenerative therapy and tissue engineering for the treatment of end-stage cardiac failure

PubMed Central

Finosh, G.T.; Jayabalan, Muthu

2012-01-01

Regeneration of myocardium through regenerative therapy and tissue engineering is appearing as a prospective treatment modality for patients with end-stage heart failure. Focusing on this area, this review highlights the new developments and challenges in the regeneration of myocardial tissue. The role of various cell sources, calcium ion and cytokine on the functional performance of regenerative therapy is discussed. The evolution of tissue engineering and the role of tissue matrix/scaffold, cell adhesion and vascularisation on tissue engineering of cardiac tissue implant are also discussed. PMID:23507781

Regenerative therapy and tissue engineering for the treatment of end-stage cardiac failure: new developments and challenges.

PubMed

Finosh, G T; Jayabalan, Muthu

2012-01-01

Regeneration of myocardium through regenerative therapy and tissue engineering is appearing as a prospective treatment modality for patients with end-stage heart failure. Focusing on this area, this review highlights the new developments and challenges in the regeneration of myocardial tissue. The role of various cell sources, calcium ion and cytokine on the functional performance of regenerative therapy is discussed. The evolution of tissue engineering and the role of tissue matrix/scaffold, cell adhesion and vascularisation on tissue engineering of cardiac tissue implant are also discussed.

Alternative Bio-Derived JP-8 Class Fuel and JP-8 Fuel: Flame Tube Combustor Test Results Compared using a GE TAPS Injector Configuration

NASA Technical Reports Server (NTRS)

Hicks, Yolanda R.; Tedder, Sarah A.; Anderson, Robert C.

2016-01-01

This paper presents results from tests in a flame tube facility, where a bio-derived alternate fuel was compared with JP-8 for emissions and general combustion performance. A research version of General Electric Aviation (GE) TAPS injector was used for the tests. Results include combustion efficiency from gaseous emission measurements, 2D planar laser-based imaging as well as basic flow visualization of the flame. Four inlet test conditions were selected that simulate various engine power conditions relevant to NASA Fundamental Aeronautics Supersonics Project and Environmentally Responsible Aviation Program. One inlet condition was a pilot-only test point. The other three inlet conditions incorporated fuel staging via a split between the pilot and main circuits of either 10%/90% or 20%/80%. For each engine power condition, three fuel mixes were used: 100% JP-8; 100% alternative; and a blend of the two, containing 75% alternative. Results for the inlet cases that have fuel split between pilot and main, indicate that fuel from the pilot appears to be evaporated by the time it reaches the dome exit. Main circuit liquid evaporates within a downstream distance equal to annulus height, no matter the fuel. Some fuel fluorescence images for a 10%/90% fuel staging case show a distinct difference between JP-8 and bio-derived fuel. OH PLIF results indicate that OH forms in a region more centrally-located for the JP-8 case downstream of the pilot, in its central recirculation region (CRZ). For the bio-derived Hydrotreated Renewable Jet (HRJ) fuel, however, we do not see much OH in the CRZ. The OH image structure near the dome exit is similar for the two fuels, but farther downstream the OH in the CRZ is much more apparent for the JP-8 than for the alternate fuel. For all conditions, there was no discernable difference between fuel types in combustion efficiency or emissions.

Increase Productivity Through Knowledge Management

NASA Astrophysics Data System (ADS)

Gavrikova, N. A.; Dolgih, I. N.; Dyrina, E. N.

2016-04-01

Increase in competition level requires companies to improve the efficiency of work force use characterized by labor productivity. Professional knowledge of staff and its experience play the key role in it. The results of Extrusion Line operator’s working time analysis are performed in this article. The analysis revealed that the reasons of working time ineffective use connected with inadequate information exchange and knowledge management in the company. Authors suggest the way to solve this problem: the main sources of knowledge in engineering enterprise have been defined, the conditions of success and the stages of knowledge management control have been stated.

Nuclear electric propulsion mission engineering study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1973-01-01

Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied. The NEP stage design provides both inherent reliability and high payload mass capability. The NEP stage and payload integration was found to be compatible with the space shuttle.

Saturn V First Stage (S-1C) Ready for Assembly AT KSC

NASA Technical Reports Server (NTRS)

1968-01-01

This photograph shows the Saturn V first stage (S-1C) in the Vehicle Assembly Building at Kennedy Space Center ready to be mated with the second and third stages to complete the assembly of a Saturn V launch vehicle. This particular Saturn V was used for Apollo 6, which was a systems test flight. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

An investigation of late-combustion soot burnout in a DI diesel engine using simultaneous planar imaging of soot and OH radical

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

John E. Dec; Peter L. Kelly-Zion

Diesel engine design continues to be driven by the need to improve performance while at the same time achieving further reductions in emissions. The development of new designs to accomplish these goals requires an understanding of how the emissions are produced in the engine. Laser-imaging diagnostics are uniquely capable of providing this information, and the understanding of diesel combustion and emissions formation has been advanced considerably in recent years by their application. However, previous studies have generally focused on the early and middle stages of diesel combustion. These previous laser-imaging studies do provide important insight into the soot formation andmore » oxidation processes during the main combustion event. They indicate that prior to the end of injection, soot formation is initiated by fuel-rich premixed combustion (equivalence ratio > 4) near the upstream limit of the luminous portion of the reacting fuel jet. The soot is then oxidized at the diffusion flame around the periphery of the luminous plume. Under typical diesel engine conditions, the diffusion flame does not burn the remaining fuel and soot as rapidly as it is supplied, resulting in an expanding region of rich combustion products and soot. This is evident in natural emission images by the increasing size of the luminous soot cloud prior to the end of injection. Hence, the amount of soot in the combustion chamber typically increases until shortly after the end of fuel injection, at which time the main soot formation period ends and the burnout phase begins. Sampling valve and two-color pyrometry data indicate that the vast majority (more than 90%) of the soot formed is oxidized before combustion ends; however, it is generally thought that a small fraction of this soot from the main combustion zones is not consumed and is the source of tail pipe soot emissions.« less

Development Status of the J-2X

NASA Technical Reports Server (NTRS)

Kynard, Mike; Vilja, John

2008-01-01

In June 2006, the NASA Marshall Space Flight Center (MSFC) and Pratt & Whitney Rocketdyne began development of an engine for use on the Ares I crew launch vehicle and the Ares V cargo launch vehicle. The development program will be completed in December 2012 at the end of a Design Certification Review and after certification testing of two flight configuration engines. A team of over 600 people within NASA and Pratt & Whitney Rocketdyne are currently working to prepare for the fall 2008 Critical Design Review (CDR), along with supporting an extensive risk mitigation test program. The J-2X will power the Ares I upper stage and the Ares V earth departure stage (EDS). The initial use will be in the Ares I, used to launch the Orion crew exploration vehicle. In this application, it will power the upper stage after being sent aloft on a Space Shuttle-derived. 5-segment solid rocket booster first stage. In this mission. the engine will ignite at altitude and provide the necessary acceleration force to allow the Orion to achieve orbital velocity. The Ares I upper stage, along with the J-2X. will then be expended. On the Ares V. first stage propulsion is provided by five RS-68B engines and two 5-segment boosters similar to the Ares I configuration. In the Ares V mission. the J-2X is first started to power the EDS and its payload. the Altair lunar lander. into earth orbit, then shut-down and get prepared for its next start. The EDS/Altair will remain in a parking orbit, awaiting rendezvous and docking with Orion. Once the two spacecraft are mated, the J-2X will be restarted to achieve earth departure velocity. After powering the Orion and Altair, the EDS will be expended. By using the J-2X Engine in both applications, a significant infrastructure cost savings is realized. Only one engine development is required, and the sustaining engineering and flight support infrastructures can be combined. There is also flexibility for changing, the production and flight manifest because a single production line can support both missions with minimal differences between each engine configuration kit.

Saturn Apollo Program

NASA Image and Video Library

1962-11-16

The Saturn I (SA-3) flight lifted off from Kennedy Space Center launch Complex 34, November 16, 1962. The third launch of Saturn launch vehicles, developed at the Marshall Space Flight Center (MSFC) under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet. and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. During the SA-3 flight, the upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. The water was released at an altitude of 65 miles, where within only 5 seconds, it expanded into a massive ice cloud 4.6 miles in diameter. Release of this vast quantity of water in a near-space environment marked the first purely scientific large-scale experiment.

n/a

NASA Image and Video Library

1963-03-28

The Saturn I (SA-4) flight lifted off from Kennedy Space Center launch Complex 34, March 28, 1963. The fourth launch of Saturn launch vehicles developed at the Marshall Space Flight Center (MSFC), under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. Like SA-3, the SA-4 flight’s upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. Release of this vast quantity of water in a near-space environment marked the second purely scientific large-scale experiment. The SA-4 was the last Block I rocket launch.

n/a

NASA Image and Video Library

1963-03-28

The Saturn I (SA-4) flight lifted off from Kennedy Space Center launch Complex 34, March 28, 1963. The fourth launch of Saturn launch vehicles, developed at the Marshall Space Flight Center (MSFC) under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. Like SA-3, the SA-4 flight’s upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. Release of this vast quantity of water in a near-space environment marked the second purely scientific large-scale experiment. The SA-4 was the last Block I rocket launch.

Saturn Apollo Program

NASA Image and Video Library

1965-01-01

The Saturn V first stages were test fired at the Mississippi Test Facility and at the Marshall Space Flight Center (MSFC). Five F-1 engines powered the first stage, each developing 1.5 million pounds of thrust. The first stage, known as the S-IC stage, burned over 15 tons of propellant per second during its 2.5 minutes of operation to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour. The stage was 138 feet long and 33 feet in diameter. This photograph shows the test firing of an F-1 engine at the MSFC's S-IC Static Test Firing Facility.

A review of findings of a study of rocket based combined cycle engines applied to extensively axisymmetric single stage to orbit vehicles

NASA Technical Reports Server (NTRS)

Foster, Richard W.

1992-01-01

Extensively axisymmetric and non-axisymmetric Single Stage To Orbit (SSTO) vehicles are considered. The information is presented in viewgraph form and the following topics are presented: payload comparisons; payload as a percent of dry weight - a system hardware cost indicator; life cycle cost estimations; operations and support costs estimation; selected engine type; and rocket engine specific impulse calculation.

Saturn Apollo Program

NASA Image and Video Library

1966-09-15

This vintage photograph shows the 138-foot long first stage of the Saturn V being lowered to the ground following a successful static test firing at Marshall Space flight Center's S-1C test stand. The firing provided NASA engineers information on the booster's systems. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Engine systems analysis results of the Space Shuttle Main Engine redesigned powerhead initial engine level testing

NASA Technical Reports Server (NTRS)

Sander, Erik J.; Gosdin, Dennis R.

1992-01-01

Engineers regularly analyze SSME ground test and flight data with respect to engine systems performance. Recently, a redesigned SSME powerhead was introduced to engine-level testing in part to increase engine operational margins through optimization of the engine internal environment. This paper presents an overview of the MSFC personnel engine systems analysis results and conclusions reached from initial engine level testing of the redesigned powerhead, and further redesigns incorporated to eliminate accelerated main injector baffle and main combustion chamber hot gas wall degradation. The conclusions are drawn from instrumented engine ground test data and hardware integrity analysis reports and address initial engine test results with respect to the apparent design change effects on engine system and component operation.

28. Main engine air pump located to port side of ...

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

28. Main engine air pump located to port side of main engine cylinder beside engine bed. Dynamo lies aft of air pump (at right), pipe at extreme left of image carries lake water to condenser valves. - Steamboat TICONDEROGA, Shelburne Museum Route 7, Shelburne, Chittenden County, VT

HESTIA Commodities Exchange Pallet and Sounding Rocket Test Stand

NASA Technical Reports Server (NTRS)

Chaparro, Javier

2013-01-01

During my Spring 2016 internship, my two major contributions were the design of the Commodities Exchange Pallet and the design of a test stand for a 100 pounds-thrust sounding rocket. The Commodities Exchange Pallet is a prototype developed for the Human Exploration Spacecraft Testbed for Integration and Advancement (HESTIA) program. Under the HESTIA initiative the Commodities Exchange Pallet was developed as a method for demonstrating multi-system integration thru the transportation of In-Situ Resource Utilization produced oxygen and water to a human habitat. Ultimately, this prototype's performance will allow for future evaluation of integration, which may lead to the development of a flight capable pallet for future deep-space exploration missions. For HESTIA, my main task was to design the Commodities Exchange Pallet system to be used for completing an integration demonstration. Under the guidance of my mentor, I designed, both, the structural frame and fluid delivery system for the commodities pallet. The fluid delivery system includes a liquid-oxygen to gaseous-oxygen system, a water delivery system, and a carbon-dioxide compressors system. The structural frame is designed to meet safety and transportation requirements, as well as the ability to interface with the ER division's Portable Utility Pallet. The commodities pallet structure also includes independent instrumentation oxygen/water panels for operation and system monitoring. My major accomplishments for the commodities exchange pallet were the completion of the fluid delivery systems and the structural frame designs. In addition, parts selection was completed in order to expedite construction of the prototype, scheduled to begin in May of 2016. Once the commodities pallet is assembled and tested it is expected to complete a fully integrated transfer demonstration with the ISRU unit and the Environmental Control and Life Support System test chamber in September of 2016. In addition to the development of the Commodities Exchange Pallet, I also assisted in preparation for testing the upper stage of a sounding rocket developed as a Center Innovation Fund project. The main objective of this project is to demonstrate the integration between a propulsion system and a solid oxide fuel cell (SOFC). The upper stage and SOFC are scheduled to complete an integrated test in August of 2016. As part of preparation for scheduled testing, I was responsible for designing the upper stage's test stand/support structure and main engine plume deflector to be used during hot-fire testing (fig. 3). The structural components of the test stand need to meet safety requirements for operation of the propulsion system, which consist of a 100 pounds-thrust main engine and two 15 pounds-thrust reaction control thrusters. My main accomplishment for this project was the completion of the design and the parts selection for construction of the structure, scheduled to begin late April of 2016.

Building a new predictor for multiple linear regression technique-based corrective maintenance turnaround time.

PubMed

Cruz, Antonio M; Barr, Cameron; Puñales-Pozo, Elsa

2008-01-01

This research's main goals were to build a predictor for a turnaround time (TAT) indicator for estimating its values and use a numerical clustering technique for finding possible causes of undesirable TAT values. The following stages were used: domain understanding, data characterisation and sample reduction and insight characterisation. Building the TAT indicator multiple linear regression predictor and clustering techniques were used for improving corrective maintenance task efficiency in a clinical engineering department (CED). The indicator being studied was turnaround time (TAT). Multiple linear regression was used for building a predictive TAT value model. The variables contributing to such model were clinical engineering department response time (CE(rt), 0.415 positive coefficient), stock service response time (Stock(rt), 0.734 positive coefficient), priority level (0.21 positive coefficient) and service time (0.06 positive coefficient). The regression process showed heavy reliance on Stock(rt), CE(rt) and priority, in that order. Clustering techniques revealed the main causes of high TAT values. This examination has provided a means for analysing current technical service quality and effectiveness. In doing so, it has demonstrated a process for identifying areas and methods of improvement and a model against which to analyse these methods' effectiveness.

Infectious Disease Issues in Xenotransplantation

PubMed Central

Boneva, Roumiana S.; Folks, Thomas M.; Chapman, Louisa E.

2001-01-01

Xenotransplantation, the transplantation of living organs, tissues, or cells from one species to another, is viewed as a potential solution to the existing shortage of human organs for transplantation. While whole-organ xenotransplantation is still in the preclinical stage, cellular xenotransplantation and extracorporeal perfusion applications are showing promise in early clinical trials. Advances in immunosuppressive therapy, gene engineering, and cloning of animals bring a broader array of xenotransplantation protocols closer to clinical trials. Despite several potential advantages over allotransplantation, xenotransplantation encompasses a number of problems. Immunologic rejection remains the primary hindrance. The potential to introduce infections across species barriers, another major concern, is the main focus of this review. Nonhuman primates are unlikely to be a main source for xenotransplantation products despite their phylogenetic proximity to humans. Genetically engineered pigs, bred under special conditions, are currently envisaged as the major source. Thus far, there has been no evidence for human infections caused by pig xenotransplantation products. However, the existence of xenotropic endogenous retroviruses and the clinical evidence of long-lasting porcine cell microchimerism indicate the potential for xenogeneic infections. Thus, further trials should continue under regulatory oversight, with close clinical and laboratory monitoring for potential xenogeneic infections. PMID:11148000

75 FR 27411 - Airworthiness Directives; Turbomeca Arriel 1B, 1D, 1D1, and 1S1 Turboshaft Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-05-17

.... That AD requires initial and repetitive relative position checks of the gas generator 2nd stage turbine... repetitive replacements of 2nd stage turbines on Arriel 1B, 1D, and 1D1 engines. This AD requires lowering the initial and repetitive thresholds for replacement of 2nd stage turbines on Arriel 1B, 1D, and 1D1...

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

Workmen secure a J-2 engine onto the S-IVB (second) stage thrust structure. As part of Marshall Space Center's "building block" approach to the Saturn development, the S-IVB was utilized in the Saturn IBC launch vehicle as a second stage and the Saturn V launch vehicle as a third stage. The booster, built for NASA by McDornell Douglas Corporation, was powered by a single J-2 engine, initially capable of 200,000 pounds of thrust.

Altitude Testing of Large Liquid Propellant Engines

NASA Technical Reports Server (NTRS)

Maynard, Bryon T.; Raines, Nickey G.

2010-01-01

The National Aeronautics and Space Administration entered a new age on January 14, 2004 with President Bush s announcement of the creation the Vision for Space Exploration that will take mankind back to the Moon and on beyond to Mars. In January, 2006, after two years of hard, dedicated labor, engineers within NASA and its contractor workforce decided that the J2X rocket, based on the heritage of the Apollo J2 engine, would be the new engine for the NASA Constellation Ares upper stage vehicle. This engine and vehicle combination would provide assured access to the International Space Station to replace that role played by the Space Shuttle and additionally, would serve as the Earth Departure Stage, to push the Crew Excursion Vehicle out of Earth Orbit and head it on a path for rendezvous with the Moon. Test as you fly, fly as you test was chosen to be the guiding philosophy and a pre-requisite for the engine design, development, test and evaluation program. An exhaustive survey of national test facility assets proved the required capability to test the J2X engine at high altitude for long durations did not exist so therefore, a high altitude/near space environment testing capability would have to be developed. After several agency concepts the A3 High Altitude Testing Facility proposal was selected by the J2X engine program on March 2, 2007 and later confirmed by a broad panel of NASA senior leadership in May 2007. This facility is to be built at NASA s John C. Stennis Space Center located near Gulfport, Mississippi. 30 plus years of Space Shuttle Main Engine development and flight certification testing makes Stennis uniquely suited to support the Vision For Space Exploration Return to the Moon. Propellant handling infrastructure, engine assembly facilities, a trained and dedicated workforce and a broad and varied technical support base will all ensure that the A3 facility will be built on time to support the schedule needs of the J2X engine and the ultimate flight of the first Ares I vehicle. The A3 facility will be able to simulate pre-ignition altitude from sea-level to 100,000 feet and maintain it up to 650 seconds. Additionally the facility will be able to accommodate initial ignition, shutdown and then restart test profiles. A3 will produce up to 5000 lbm/sec of superheated steam utilizing a Chemical Steam generation system. Two separate inline steam ejectors will be used to produce a test cell vacuum to simulate the 100,000 ft required altitude. Operational capability will ensure that the facility can start up and shutdown without producing adverse pressure gradients across the J2X nozzle. The facility will have a modern thrust measurement system for accurate determination of engine performance. The latest advances in data acquisition and control will be incorporated to measure performance parameters during hotfire testing. Provisions are being made in the initial design of the new altitude facility to allow for testing of other, larger engines and potential upper stage launch vehicles that might require vacuum start testing of the engines. The new facility at Stennis Space Center will be complete and ready for hotfire operations in late 2010.

Survey on effect of surface winds on aircraft design and operation and recommendations for needed wind research

NASA Technical Reports Server (NTRS)

Houbolt, J. C.

1973-01-01

A survey of the effect of environmental surface winds and gusts on aircraft design and operation is presented. A listing of the very large number of problems that are encountered is given. Attention is called to the many studies that have been made on surface winds and gusts, but development in the engineering application of these results to aeronautical problems is pointed out to be still in the embryonic stage. Control of the aircraft is of paramount concern. Mathematical models and their application in simulation studies of airplane operation and control are discussed, and an attempt is made to identify their main gaps or deficiencies. Key reference material is cited. The need for better exchange between the meteorologist and the aeronautical engineer is discussed. Suggestions for improvements in the wind and gust models are made.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians mount NASA's Soil Moisture Active Passive, or SMAP, spacecraft on a work platform. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians rotate NASA's Soil Moisture Active Passive, or SMAP, spacecraft to begin processing. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians have rotated NASA's Soil Moisture Active Passive, or SMAP, spacecraft to begin processing. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians inspect NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians begin processing of NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, an engineer inspects NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians remove a protective covering from NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians prepare a component of NASA's Soil Moisture Active Passive, or SMAP, spacecraft for a lift by a crane. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians rotate NASA's Soil Moisture Active Passive, or SMAP, spacecraft to begin processing. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians use a crane to move NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians have rotated NASA's Soil Moisture Active Passive, or SMAP, spacecraft to begin processing. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Spacecraft Rotate & Placed on Fixture

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians prepare a component of NASA's Soil Moisture Active Passive, or SMAP, spacecraft for a lift by a crane. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians remove a protective covering from NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians inspect NASA's Soil Moisture Active Passive, or SMAP, spacecraft. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

SMAP Lift to CR

NASA Image and Video Library

2014-10-16

Inside the Astrotech payload processing facility on Vandenberg Air Force Base in California, engineers and technicians use a crane to move a component of NASA's Soil Moisture Active Passive, or SMAP, spacecraft for a lift by a crane. SMAP will launch on a Delta II 7320 configuration vehicle featuring a United Launch Alliance first stage booster powered by an Aerojet Rocketdyne RS-27A main engine and three Alliant Techsystems, or ATK, strap-on solid rocket motors. Once on station in Earth orbit, SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch from Space Launch Complex 2 is targeted for Jan. 29, 2015.

Numerical simulation of interaction between chemically active exhaust and a jet blast deflector

NASA Astrophysics Data System (ADS)

Korotaeva, T. A.; Turchinovich, A. O.

2017-10-01

The interaction of chemically active exhausts of aircraft engines with jet blast deflector (JBD) of various configurations has been considered at the stage of ground run procedure. The problem is modeled in the 3-D approximation in the framework of the numerical solution of the Navier-Stokes equations taking into account the kinetic model of the interaction of between the components of engine exhaust and air. A complex field of gasdynamic flow that is realized when jets emerge from nozzles and interact with each other, with air, with a gas deflector has been studied. The main purpose of the study is to prove the concept that it is possible to generate a vortex flow that can not only change the direction of the jets, but also contribute to the lifting of the mass of pollutants and their dispersion in the atmosphere using a gas deflector shape.

3-D thermal analysis using finite difference technique with finite element model for improved design of components of rocket engine turbomachines for Space Shuttle Main Engine SSME

NASA Technical Reports Server (NTRS)

Sohn, Kiho D.; Ip, Shek-Se P.

1988-01-01

Three-dimensional finite element models were generated and transferred into three-dimensional finite difference models to perform transient thermal analyses for the SSME high pressure fuel turbopump's first stage nozzles and rotor blades. STANCOOL was chosen to calculate the heat transfer characteristics (HTCs) around the airfoils, and endwall effects were included at the intersections of the airfoils and platforms for the steady-state boundary conditions. Free and forced convection due to rotation effects were also considered in hollow cores. Transient HTCs were calculated by taking ratios of the steady-state values based on the flow rates and fluid properties calculated at each time slice. Results are presented for both transient plots and three-dimensional color contour isotherm plots; they were also converted into universal files to be used for FEM stress analyses.

A new collection of real world applications of fractional calculus in science and engineering

NASA Astrophysics Data System (ADS)

Sun, HongGuang; Zhang, Yong; Baleanu, Dumitru; Chen, Wen; Chen, YangQuan

2018-11-01

Fractional calculus is at this stage an arena where many models are still to be introduced, discussed and applied to real world applications in many branches of science and engineering where nonlocality plays a crucial role. Although researchers have already reported many excellent results in several seminal monographs and review articles, there are still a large number of non-local phenomena unexplored and waiting to be discovered. Therefore, year by year, we can discover new aspects of the fractional modeling and applications. This review article aims to present some short summaries written by distinguished researchers in the field of fractional calculus. We believe this incomplete, but important, information will guide young researchers and help newcomers to see some of the main real-world applications and gain an understanding of this powerful mathematical tool. We expect this collection will also benefit our community.

STS-41 Ulysses Breakfast, Suit-up, C-7 Exit, Launch and ISOS Cam Views

NASA Technical Reports Server (NTRS)

1990-01-01

Live footage shows the crewmembers of STS-41, Commander Richard N. Richards, Pilot Robert D. Cabana, Mission Specialists William M. Shepherd, Bruce E. Melnick, and Thomas D. Akers, participating in the traditional activities the day of their flight. The crew are seen eating breakfast, suiting-up, walking out to the Astronaut-Van, putting on life vests in the 'White Room' area, and entering the crew module of the Discovery Orbiter. Footage also includes preparation of the Ulysses Payload. Engineers are seen loading Ulysses to the upper stage, transferring Discovery to an upright position, bolting Discovery to the external tank, rolling Discovery out to the launch pad, and finally installing the Ulysses Payload inside Discovery. Also shown are both night and morning panoramic shots of the shuttle on the pad, main engine start, ignition, liftoff, booster separation, and various camera views of the launch.

Engine/vehicle integration for vertical takeoff and landing single stage to orbit vehicles

NASA Astrophysics Data System (ADS)

Weegar, R. K.

1992-08-01

SSTO vehicles design which is currently being developed under the Single Stage Rocket Technology program of the Strategic Defense Initiative Organization is discussed. Particular attention is given to engine optimization and integration of ascent, orbital, and landing propulsion requirements into a single system.

Compressor Modeling for Engine Control and Maintenance

DTIC Science & Technology

2011-07-01

four compressor stages, while the high pressure compressor (HPC) consists of a set of variable pitch inlet guide vanes ( IGVs ) and 12 compressor...bleed valves at stages 5, 14 and 17, along with the variable IGVs and stators within the engine, are used to relieve the pressure and prevent

Evaluating Games-Based Learning

ERIC Educational Resources Information Center

Hainey, Thomas; Connolly, Thomas

2010-01-01

A highly important part of software engineering education is requirements collection and analysis, one of the initial stages of the Software Development Lifecycle. No other conceptual work is as difficult to rectify at a later stage or as damaging to the overall system if performed incorrectly. As software engineering is a field with a reputation…

Saturn Apollo Program

NASA Image and Video Library

1964-03-03

Two technicians apply insulation to the outer surface of the S-II second stage booster for the Saturn V moon rocket. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A NASA technician is dwarfed by the gigantic Third Stage (S-IVB) as it rests on supports in a facility at KSC. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

Workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, install the last engine on the S-IB stage. Developed by the Marshall Space Flight Center (MSFC) and built by the Chrysler Corporation at MAF, the S-IB stage utilized eight H-1 engines to produce a combined thrust of 1,600,000 pounds.

Computational Analysis of Advanced Shape-Memory Alloy Devices Through a Robust Modeling Framework

NASA Astrophysics Data System (ADS)

Scalet, Giulia; Conti, Michele; Auricchio, Ferdinando

2017-06-01

Shape-memory alloys (SMA) provide significant advantages in various industrial fields, but their manufacturing and commercialization are currently hindered. This is attributed mainly to the poor knowledge of material behavior and the lack of standards in its mechanical characterization. SMA products are usually developed by trial-and-error testing to address specific design requirements, thus increasing costs and time. The development of simulation tools offers a possible solution to assist engineers and designers and allows to better understand SMA transformation phenomena. Accordingly, the purpose of the present paper is to numerically analyze and predict the response of spring-like actuators and septal occluders, which are industrial components exploiting the shape-memory and pseudoelastic properties of SMAs, respectively. The methodology includes two main stages: the implementation of the three-dimensional phenomenological model known as Souza- Auricchio model and the finite element modeling of the device. A discussion about the steps of each stage, as parameter identification and model generalizations, is provided. Validation results are presented through a comparison with the results of a performed experimental campaign. The framework proves good prediction capabilities and allows to reduce the number of experimental tests in the future.

An overview of the current technology relevant to the design and development of the Space Transportation Main Engine (STME)

NASA Technical Reports Server (NTRS)

Das, Digendra K.

1991-01-01

The objective of this project was to review the latest literature relevant to the Space Transportation Main Engine (STME). The search was focused on the following engine components: (1) gas generator; (2) hydrostatic/fluid bearings; (3) seals/clearances; (4) heat exchanges; (5) nozzles; (6) nozzle/main combustion chamber joint; (7) main injector face plate; and (8) rocket engine.

Preburner of Staged Combustion Rocket Engine

NASA Technical Reports Server (NTRS)

Yost, M. C.

1978-01-01

A regeneratively cooled LOX/hydrogen staged combustion assembly system with a 400:1 expansion area ratio nozzle utilizing an 89,000 Newton (20,000 pound) thrust regeneratively cooled thrust chamber and 175:1 tubular nozzle was analyzed, assembled, and tested. The components for this assembly include two spark/torch oxygen-hydrogen igniters, two servo-controlled LOX valves, a preburner injector, a preburner combustor, a main propellant injector, a regeneratively cooled combustion chamber, a regeneratively cooled tubular nozzle with an expansion area ratio of 175:1, an uncooled heavy-wall steel nozzle with an expansion area ratio of 400:1, and interconnecting ducting. The analytical effort was performed to optimize the thermal and structural characteristics of each of the new components and the ducting, and to reverify the capabilities of the previously fabricated components. The testing effort provided a demonstration of the preburner/combustor chamber operation, chamber combustion efficiency and stability, and chamber and nozzle heat transfer.

Stochastic sensitivity measure for mistuned high-performance turbines

NASA Technical Reports Server (NTRS)

Murthy, Durbha V.; Pierre, Christophe

1992-01-01

A stochastic measure of sensitivity is developed in order to predict the effects of small random blade mistuning on the dynamic aeroelastic response of turbomachinery blade assemblies. This sensitivity measure is based solely on the nominal system design (i.e., on tuned system information), which makes it extremely easy and inexpensive to calculate. The measure has the potential to become a valuable design tool that will enable designers to evaluate mistuning effects at a preliminary design stage and thus assess the need for a full mistuned rotor analysis. The predictive capability of the sensitivity measure is illustrated by examining the effects of mistuning on the aeroelastic modes of the first stage of the oxidizer turbopump in the Space Shuttle Main Engine. Results from a full analysis mistuned systems confirm that the simple stochastic sensitivity measure predicts consistently the drastic changes due to misturning and the localization of aeroelastic vibration to a few blades.

Saturn Apollo Program

NASA Image and Video Library

1961-05-16

On October 27, 1961, the Marshall Space Flight Center (MSFC) and the Nation marked a high point in the 3-year-old Saturn development program when the first Saturn vehicle flew a flawless 215-mile ballistic trajectory from Cape Canaveral, Florida. SA-1 is pictured here, five months before launch, in the MSFC test stand on May 16, 1961. Developed and tested at MSFC under the direction of Dr. Wernher von Braun, SA-1 incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet. and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks, as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle.

Advanced space engine preliminary design. [liquid hydrogen/liquid oxygen upper stage engine for space tug application

NASA Technical Reports Server (NTRS)

Zachary, A. T.

1973-01-01

Analysis and design of an optimum LO2/LH2, combustion topping cycle, 88,964 Newtons (20,000-pound) thrust, liquid rocket engine was conducted. The design selected is well suited to high-energy, upper-stage engine applications such as the Space Tug and embodies features directed toward optimization of vehicle performance. A configuration selection was conducted based on prior Air Force Contracts, and additional criteria for optimum stage performance. Following configuration selection, analyses and design of the major components and engine systems were conducted to sufficient depth to provide layout drawings suitable for subsequent detailing. In addition, engine packaging to a common interface and a retractable nozzle concept were defined. Alternative development plans and related costs were also established. The design embodies high-performance, low-weight, low NPSH requirements (saturated propellant inlet conditions at start), idle-mode operation, and autogenous pressurization. The design is the result of the significant past and current LO2/LH2 technology efforts of the NASA centers and the Air Force, as well as company-funded programs.

SLS Engine Section Test Article Moves From NASA Barge Pegasus To Test Stand at NASA’s Marshall Space Flight Center

NASA Image and Video Library

2017-05-18

The NASA barge Pegasus made its first trip to NASA’s Marshall Space Flight Center in Huntsville, Alabama on May 15. It arrived carrying the first piece of Space Launch System hardware built at NASA's Michoud Assembly Facility in New Orleans. The barge left Michoud on April 28 with the core stage engine section test article, traveling 1,240 miles by river to Marshall. The rocket's engine section is the bottom of the core stage and houses the four RS-25 engines. The engine section test article was moved from the barge to Marshall’s Building 4619 where it will be tested. The bottom part of the test article is structurally the same as the engine section that will be flown as part of the SLS core stage. The shiny metal top part simulates the rocket's liquid hydrogen tank, which is the fuel tank that joins to the engine section. The test article will endure tests that pull, push, and bend it, subjecting it to millions of pounds of force. This ensures the structure can withstand the incredible stresses produced by the 8.8 million pounds of thrust during launch and ascent.

SLS Engine Section Test Article Arrives at Marshall on NASA Barge Pegasus

NASA Image and Video Library

2017-05-16

The NASA barge Pegasus made it’s first trip to NASA’s Marshall Space Flight Center in Huntsville, Alabama on May 15. It arrived carrying the first piece of Space Launch System hardware built at NASA's Michoud Assembly Facility in New Orleans. The barge left Michoud on April 28 with the core stage engine section test article, traveling 1,240 miles by river to Marshall. The rocket's engine section is the bottom of the core stage and houses the four RS-25 engines. The engine section test article will be moved to Marshall’s Building 4619 where it will be tested. The bottom part of the test article is structurally the same as the engine section that will be flown as part of the SLS core stage. The shiny metal top part simulates the rocket's liquid hydrogen tank, which is the fuel tank that joins to the engine section. The test article will endure tests that pull, push, and bend it, subjecting it to millions of pounds of force. This ensures the structure can withstand the incredible stresses produced by the 8.8 million pounds of thrust during launch and ascent.

KSC-07pd1271

NASA Image and Video Library

2007-05-24

KENNEDY SPACE CENTER, FLA. -- In Space Shuttle Maine Engine Shop, workers get ready to install an engine controller in one of the three main engines (behind them) of the orbiter Discovery. The controller is an electronics package mounted on each space shuttle main engine. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings. Discovery is the designated orbiter for mission STS-120 to the International Space Station. It will carry a payload that includes the Node 2 module, named Harmony. Launch is targeted for no earlier than Oct. 20. Photo credit: NASA/Cory Huston

KSC-07pd1272

NASA Image and Video Library

2007-05-24

KENNEDY SPACE CENTER, FLA. -- In the Space Shuttle Maine Engine Shop, workers are installing an engine controller in one of the three main engines of the orbiter Discovery. The controller is an electronics package mounted on each space shuttle main engine. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings. Discovery is the designated orbiter for mission STS-120 to the International Space Station. It will carry a payload that includes the Node 2 module, named Harmony. Launch is targeted for no earlier than Oct. 20. Photo credit: NASA/Cory Huston

KSC-07pd1274

NASA Image and Video Library

2007-05-24

KENNEDY SPACE CENTER, FLA. -- In the Space Shuttle Maine Engine Shop, workers check the installation of an engine controller in one of the three main engines of the orbiter Discovery. The controller is an electronics package mounted on each space shuttle main engine. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings. Discovery is the designated orbiter for mission STS-120 to the International Space Station. It will carry a payload that includes the Node 2 module, named Harmony. Launch is targeted for no earlier than Oct. 20. Photo credit: NASA/Cory Huston

KSC-07pd1273

NASA Image and Video Library

2007-05-24

KENNEDY SPACE CENTER, FLA. -- In the Space Shuttle Maine Engine Shop, workers are installing an engine controller in one of the three main engines of the orbiter Discovery. The controller is an electronics package mounted on each space shuttle main engine. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings. Discovery is the designated orbiter for mission STS-120 to the International Space Station. It will carry a payload that includes the Node 2 module, named Harmony. Launch is targeted for no earlier than Oct. 20. Photo credit: NASA/Cory Huston

KSC-07pd1270

NASA Image and Video Library

2007-05-24

KENNEDY SPACE CENTER, FLA. -- In the Space Shuttle Maine Engine Shop, workers get ready to install an engine controller in one of the three main engines of the orbiter Discovery. The controller is an electronics package mounted on each space shuttle main engine. It contains two digital computers and the associated electronics to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings. Discovery is the designated orbiter for mission STS-120 to the International Space Station. It will carry a payload that includes the Node 2 module, named Harmony. Launch is targeted for no earlier than Oct. 20. Photo credit: NASA/Cory Huston

NASA's Space Launch System Takes Shape

NASA Technical Reports Server (NTRS)

Askins, Bruce; Robinson, Kimberly F.

2017-01-01

Major hardware and software for NASA's Space Launch System (SLS) began rolling off assembly lines in 2016, setting the stage for critical testing in 2017 and the launch of a major new capability for deep space human exploration. SLS continues to pursue a 2018 first launch of Exploration Mission 1 (EM-1). At NASA's Michoud Assembly Facility near New Orleans, LA, Boeing completed welding of structural test and flight liquid hydrogen tanks, and engine sections. Test stands for core stage structural tests at NASA's Marshall Space Flight Center, Huntsville, AL. neared completion. The B2 test stand at NASA's Stennis Space Center, MS, completed major structural renovation to support core stage green run testing in 2018. Orbital ATK successfully test fired its second qualification solid rocket motor in the Utah desert and began casting the motor segments for EM-1. Aerojet Rocketdyne completed its series of test firings to adapt the heritage RS-25 engine to SLS performance requirements. Production is under way on the first five new engine controllers. NASA also signed a contract with Aerojet Rocketdyne for propulsion of the RL10 engines for the Exploration Upper Stage. United Launch Alliance delivered the structural test article for the Interim Cryogenic Propulsion Stage to MSFC for tests and construction was under way on the flight stage. Flight software testing at MSFC, including power quality and command and data handling, was completed. Substantial progress is planned for 2017. Liquid oxygen tank production will be completed at Michoud. Structural testing at Marshall will get under way. RS-25 hotfire testing will verify the new engine controllers. Core stage horizontal integration will begin. The core stage pathfinder mockup will arrive at the B2 test stand for fit checks and tests. EUS will complete preliminary design review. This paper will discuss the technical and programmatic successes and challenges of 2016 and look ahead to plans for 2017.

Analysis of turbojet combustion chamber performances based on flow field simplified mathematical model

NASA Astrophysics Data System (ADS)

Rotaru, Constantin

2017-06-01

In this paper are presented some results about the study of combustion chamber geometrical configurations that are found in aircraft gas turbine engines. The main focus of this paper consists in a study of a new configuration of the aircraft engine combustion chamber with an optimal distribution of gas velocity in front of the turbine. This constructive solution could allow a lower engine rotational speed, a lower temperature in front of the first stage of the turbine and the possibility to increase the turbine pressure ratio. The Arrhenius relationship, which describes the basic dependencies of the reaction rate on pressure, temperature and concentration has been used. and the CFD simulations were made with jet A fuel (which is presented in the Fluent software database) for an annular flame tube with 24 injectors. The temperature profile at the turbine inlet exhibits nonuniformity due to the number of fuel injectors used in the circumferential direction, the spatial nonuniformity in dilution air cooling and mixing characteristics as well as other secondary flow patterns and instabilities that are set up in the flame tube.

78 FR 35747 - Airworthiness Directives; Engine Alliance Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2013-06-14

... Airworthiness Directives; Engine Alliance Turbofan Engines AGENCY: Federal Aviation Administration (FAA), DOT... Alliance GP7270 and GP7277 turbofan engines. This AD was prompted by damage to the high-pressure compressor... Alliance GP7270 and GP7277 turbofan engines with a high-pressure compressor (HPC) stage 6 disk, part number...

Experimental measurements and analytical analysis related to gas turbine heat transfer. Part 1: Time-averaged heat-flux and surface-pressure measurements on the vanes and blades of the SSME fuel-side turbine and comparison with prediction. Part 2: Phase-resolved surface-pressure and heat-flux measurements on the first blade of the SSME fuel-side turbine

NASA Technical Reports Server (NTRS)

1994-01-01

Time averaged Stanton number and surface-pressure distributions are reported for the first-stage vane row, the first stage blade row, and the second stage vane row of the Rocketdyne Space Shuttle Main Engine two-stage fuel-side turbine. Unsteady pressure envelope measurements for the first blade are also reported. These measurements were made at 10 percent, 50 percent, and 90 percent span on both the pressure and suction surfaces of the first stage components. Additional Stanton number measurements were made on the first stage blade platform blade tip, and shroud, and at 50 percent span on the second vane. A shock tube was used as a short duration source of heated and pressurized air to which the turbine was subjected. Platinum thin-film heat flux gages were used to obtain the heat flux measurements, while miniature silicon-diaphragm flush-mounted pressure transducers were used to obtain the pressure measurements. The first stage vane Stanton number distributions are compared with predictions obtained using a version of STAN5 and a quasi-3D Navier-Stokes solution. This same quasi-3D N-S code was also used to obtain predictions for the first blade and the second vane.

Closeup View of the Space Shuttle Main Engine (SSME) 2044 ...

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

Close-up View of the Space Shuttle Main Engine (SSME) 2044 mounted in a SSME Engine Handler in the SSME processing Facility at Kennedy Space Center. This view shows SSME 2044 with its expansion nozzle removed and an Engine Leak-Test Plug is set in the throat of the Main Combustion Chamber in the approximate center of the image, the insulated, High-Pressure Fuel Turbopump sits below that and the Low Pressure Oxidizer Turbopump Discharge Duct sits towards the top of the engine assembly in this view. - Space Transportation System, Space Shuttle Main Engine, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Investigation of the Effect of the Non-uniform Flow Distribution After Compressor of Gas Turbine Engine on Inlet Parameters of the Turbine

NASA Astrophysics Data System (ADS)

Orlov, M. Yu; Lukachev, S. V.; Anisimov, V. M.

2018-01-01

The position of combustion chamber between compressor and turbine and combined action of these elements imply that the working processes of all these elements are interconnected. One of the main requirements of the combustion chamber is the formation of the desirable temperature field at the turbine inlet, which can realize necessary durability of nozzle assembly and blade wheel of the first stage of high-pressure turbine. The method of integrated simulation of combustion chamber and neighboring nodes (compressor and turbine) was developed. On the first stage of the study, this method was used to investigate the influence of non-uniformity of flow distribution, occurred after compressor blades on combustion chamber workflow. The goal of the study is to assess the impact of non-uniformity of flow distribution after the compressor on the parameters before the turbine. The calculation was carried out in a transient case for some operation mode of the engine. The simulation showed that the inclusion of compressor has an effect on combustion chamber workflow and allows us to determine temperature field at the turbine inlet and assesses its durability more accurately. In addition, the simulation with turbine showed the changes in flow velocity distribution and pressure in combustion chamber.

Turbine disk cavity aerodynamics and heat transfer

NASA Technical Reports Server (NTRS)

Johnson, B. V.; Daniels, W. A.

1992-01-01

Experiments were conducted to define the nature of the aerodynamics and heat transfer for the flow within the disk cavities and blade attachments of a large-scale model, simulating the Space Shuttle Main Engine (SSME) turbopump drive turbines. These experiments of the aerodynamic driving mechanisms explored the following: (1) flow between the main gas path and the disk cavities; (2) coolant flow injected into the disk cavities; (3) coolant density; (4) leakage flows through the seal between blades; and (5) the role that each of these various flows has in determining the adiabatic recovery temperature at all of the critical locations within the cavities. The model and the test apparatus provide close geometrical and aerodynamic simulation of all the two-stage cavity flow regions for the SSME High Pressure Fuel Turbopump and the ability to simulate the sources and sinks for each cavity flow.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

This photograph shows the fuel tank assembly for the Saturn V S-IC (first) stage being transported to the Marshall Space Flight Center, building 4705 for mating to the liquid oxygen (LOX) tank. The fuel tank carried kerosene (RP-1) as its fuel. The S-IC stage used five F-1 engines, that used kerosene and liquid oxygen as propellant and each engine provided 1,500,000 pounds of thrust. This stage lifted the entire vehicle and Apollo spacecraft from the launch pad.

Numerical simulation of stress-strain state of electrophoretic shell molds

NASA Astrophysics Data System (ADS)

Sviridov, A. V.; Odinokov, V. I.; Dmitriev, E. A.; Evstigneev, A. I.; Bashkov, O. V.

2017-10-01

In the foundry engineering, castings obtained in one-piece non-gas-generating high-refractory electrophoretic shell molds (ShM) by investment patterns (IP) have an increased rejects percentage associated with low deformation resistance and crack resistance of the SM at different stages of their formation and manufacturing. Crack resistance of the ShM based on IP depends mainly on their stress-strain state (SSS) at various stages of mold forming. SSS decrease significantly improves their crack resistance and decreases their rejects percentage of castings occurring due to clogging and surface defects. In addition, the known methods of decreasing the SSS are still poorly understood. Thus, current research trends are to determine SSS at each stage of ShM forming and develop the ways to decrease it. Theoretical predicting of crack formation in multiple-layer axisymmetric shell molds is given in the work [1], and SSS of multiple-layer axisymmetric shell molds is given in the work [2]. Monolayer electrophoretic ShM had a lack of concern in this field, thus it became an argument for the present workMathematical Model of ShM SSS

General view of a Space Shuttle Main Engine (SSME) mounted ...

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

General view of a Space Shuttle Main Engine (SSME) mounted on an SSME engine handler, taken in the SSME Processing Facility at Kennedy Space Center. The most prominent features of the engine assembly in this view are the Low-Pressure Fuel Turbopump Discharge Duct looping around the right side and underneath the assembly, the High-Pressure Fuel Turbopump located on the lower left portion of the assembly, the Engine Controller and Main Fuel Valve Hydraulic Actuator located on the upper portion of the assembly and the Low-Pressure Oxidizer Turbopump Discharge Duct at the top of the engine assembly in this view. - Space Transportation System, Space Shuttle Main Engine, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Booster propulsion/vehicle impact study

NASA Technical Reports Server (NTRS)

Weldon, Vincent; Dunn, Michael; Fink, Lawrence; Phillips, Dwight; Wetzel, Eric

1988-01-01

The use of hydrogen RP-1, propane, and methane as fuels for booster engines of launch vehicles is discussed. An automated procedure for integrated launch vehicle, engine sizing, and design optimization was used to define two stage and single stage concepts for minimum dry weight. The two stage vehicles were unmanned and used a flyback booster and partially reusable orbiter. The single stage designs were fully reusable, manned flyback vehicles. Comparisons of these vehicle designs, showing the effects of using different fuels, as well as sensitivity and trending data, are presented. In addition, the automated design technique utilized for the study is described.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

At the Marshall Space Flight Center (MSFC), the fuel tank assembly for the Saturn V S-IC-T (static test stage) fuel tank assembly is mated to the liquid oxygen (LOX) tank in building 4705. This stage underwent numerous static firings at the newly-built S-IC Static Test Stand at the MSFC west test area. The S-IC (first) stage used five F-1 engines that produced a total thrust of 7,500,000 pounds as each engine produced 1,500,000 pounds of thrust. The S-IC stage lifted the Saturn V vehicle and Apollo spacecraft from the launch pad.

Saturn Apollo Program

NASA Image and Video Library

1967-09-09

This image depicts the test firing of a J-2 engine in the S-IVB Test Stand at the Marshall Space Flight Center (MSFC). The J-2, developed by Rocketdyne under the direction of MSFC, was propelled by liquid hydrogen and liquid oxygen. A single J-2 was utilized in the S-IVB stage (the second stage for the Saturn IB and third stage for the Saturn V) and in a cluster of five for the second stage (S-II) of the Saturn V. Initially rated at 200,000 pounds of thrust, the engine was later upgraded in the Saturn V program to 230,000 pounds.

Myocardial Tissue Engineering for Regenerative Applications.

PubMed

Fujita, Buntaro; Zimmermann, Wolfram-Hubertus

2017-09-01

This review provides an overview of the current state of tissue-engineered heart repair with a special focus on the anticipated modes of action of tissue-engineered therapy candidates and particular implications as to transplant immunology. Myocardial tissue engineering technologies have made tremendous advances in recent years. Numerous different strategies are under investigation and have reached different stages on their way to clinical translation. Studies in animal models demonstrated that heart repair requires either remuscularization by delivery of bona fide cardiomyocytes or paracrine support for the activation of endogenous repair mechanisms. Tissue engineering approaches result in enhanced cardiomyocyte retention and sustained remuscularization, but may also be explored for targeted paracrine or mechanical support. Some of the more advanced tissue engineering approaches are already tested clinically; others are at late stages of pre-clinical development. Process optimization towards cGMP compatibility and clinical scalability of contractile engineered human myocardium is an essential step towards clinical translation. Long-term allograft retention can be achieved under immune suppression. HLA matching may be an option to enhance graft retention and reduce the need for comprehensive immune suppression. Tissue-engineered heart repair is entering the clinical stage of the translational pipeline. Like in any effective therapy, side effects must be anticipated and carefully controlled. Allograft implantation under immune suppression is the most likely clinical scenario. Strategies to overcome transplant rejection are evolving and may further boost the clinical acceptance of tissue-engineered heart repair.

Rocketdyne - Lunar Ascent Engine. Chapter 7, Appendix I

NASA Technical Reports Server (NTRS)

Harmon, Tim

2009-01-01

The ascent engine was the last one from the moon, and I want to focus on the idea of redundancy and teams in regard to the engine. By teams, I mean teamwork - not just within Rocketdyne. It was teamwork within Rocketdyne; it was teamwork within Grumman; it was teamwork within NASA. These were all important elements leading to the successful development of the lunar excursion module (LEM) engine. Communication, rapid response, and cooperation were all important. Another aspect that went into the development of the ascent engine was the integration of technology and of lessons learned. We pushed all the above, plus technology and lessons learned, into a program, and that led to a successful result. One of the things that I like to think about - again in retrospect - is how it is very "in" now to have integrated product and process teams. These are buzzwords for teamwork in all program phases. That s where you combine a lot of groups into a single organization to get a job done. The ascent engine program epitomized that kind of integration and focus, and because this was the mid- to late-1960s; this was new ground for Rocketdyne, Grumman, and NASA. Redundancy was really a major hallmark of the Apollo Program. Everything was redundant. Once you got the rocket going, you could even lose one of the big F-1 engines, and it would still make it to orbit. And once the first stage separated from the rest of the vehicle, the second stage could do without an engine and still make a mission. This redundancy was demonstrated when an early Apollo launch shut down a J-2 second-stage engine. Actually, they shut down two J-2 engines on that flight. Even the third stage, with its single J-2 engine, was backed up because the first two stages could toss it into a recoverable orbit. If the third stage didn't work, you were circling the earth, and you had time to recover the command module and crew. Remember how on the Apollo 13 flight, there was sufficient system redundancy even when we lost the service module. That was a magnificent effort. TRW Inc. really ought to be proud of their engine for that. (See Slide 2, Appendix I) We had planned for redundancy; we had landed on the moon. However, weight restrictions in the architecture said, "You can t have redundancy for ascent from the moon. You've got one engine. It s got to work. There is no second chance. If that ascent engine doesn't work, you re stuck there." It would not have looked good for NASA. It wouldn't have looked good for the country. There was a letter written that President Richard Nixon would read if the astronauts got stuck on the moon, expressing how sorry we were and so forth. It was a scary letter, really. The ascent engine was an engine that had to work. (See Slide 3, Appendix I).

Single-Stage, 3.4:1-Pressure-Ratio Aspirated Fan Developed and Demonstrated

NASA Technical Reports Server (NTRS)

Braunscheidel, Edward P.

2004-01-01

Researchers are constantly pursuing technologies that will increase the performance of gas turbine engines. The aspirated compressor concept discussed here would allow the compression system to perform its task with about one-half of the compressor blades. To accomplish this, the researchers applied boundary layer control to the blades, casing, and hub. This method of boundary layer control consisted of removing small amounts of air from the main flow path at critical areas of the compressor. This bleed air could be used by other systems such as engine cooling or could be re-injected into lower pressure areas that require air for enhanced performance. This effort was initiated by the Massachusetts Institute of Technology (MIT) in response to a solicitation from the Defense Advanced Research Projects Agency (DARPA) who sought to advance research in flow control technology. The NASA Glenn Research Center partnered with MIT (principal investigator), Honeywell Aircraft Engines (cycle analysis, structural analysis, and mechanical design), and Pratt & Whitney (cycle analysis and aero-analysis) to conceptualize, design, analyze, build, and test the aspirated fan stage. The aero-design and aero-analysis of this fan stage were jointly executed by MIT and Glenn to minimize the amount of bleed flow needed and to maintain the highest efficiency possible (ref. 1). Mechanical design issues were complicated by the need to have a shrouded rotor with hollow blades, with rotor stress levels beyond the capabilities of titanium. The high stress issues were addressed by designing a shroud that was filament wound with a carbon fiber/epoxy matrix, resulting in an assembly that was strong enough to handle the high stresses. Both the rotor (preceding photographs) and stator (following photograph) were fabricated in two halves and then bolted together at the hub and tip, permitting the bleed passages to be machined into each half before assembly.

Oxidizer heat exchanger component testing

NASA Technical Reports Server (NTRS)

Kmiec, T.; Kanic, P.

1986-01-01

As part of the RL10 Rocket Engine Product Improvement Program, Oxidizer Heat Exchanger (OHE) stages 1, 2, and 3 were designed and fabricated during late 1983 and early 1984. The purpose of the OHE is to provide gaseous oxygen to the propellant injector for stable engine operation at tank head idle and pumped idle operating modes. This report summarizes the OHE stages 1 and 3 rig testing, and includes the separation of the stage 1-and-2 assembly and the remanifolding of stage 1. The OHE performance analysis and analytical model modifications for both stages are also presented. The flow tests were accomplished during the time period from 9 October 1984 to 12 November 1984.

Analytical Approach for Estimating Preliminary Mass of ARES I Crew Launch Vehicle Upper Stage Structural Components

NASA Technical Reports Server (NTRS)

Aggarwal, Pravin

2007-01-01

In January 2004, President Bush gave the National Aeronautics and Space Administration (NASA) a vision for Space Exploration by setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. In response to this vision, NASA started the Constellation Program, which is a new exploration launch vehicle program. The primary mission for the Constellation Program is to carry out a series of human expeditions ranging from Low Earth Orbit to the surface of Mars and beyond for the purposes of conducting human exploration of space, as specified by the Vision for Space Exploration (VSE). The intent is that the information and technology developed by this program will provide the foundation for broader exploration activities as our operational experience grows. The ARES I Crew Launch Vehicle (CLV) has been designated as the launch vehicle that will be developed as a "first step" to facilitate the aforementioned human expeditions. The CLV Project is broken into four major elements: First Stage, Upper Stage Engine, Upper Stage (US), and the Crew Exploration Vehicle (CEV). NASA's Marshall Space Flight Center (MSFC) is responsible for the design of the CLV and has the prime responsibility to design the upper stage of the vehicle. The US is the second propulsive stage of the CLV and provides CEV insertion into low Earth orbit (LEO) after separation from the First Stage of the Crew Launch Vehicle. The fully integrated Upper Stage is a mix of modified existing heritage hardware (J-2X Engine) and new development (primary structure, subsystems, and avionics). The Upper Stage assembly is a structurally stabilized cylindrical structure, which is powered by a single J-2X engine which is developed as a separate Element of the CLV. The primary structure includes the load bearing liquid hydrogen (LH2) and liquid oxygen (LOX) propellant tanks, a Forward Skirt, the Intertank structure, the Aft Skirt and the Thrust Structure. A Systems Tunnel, which carries fluid and electrical power functions to other Elements of the CLV, is included as secondary structure. The MSFC has an overall responsibility for the integrated US element as well as structural design an thermal control of the fuel tanks, intertank, interstage, avionics, main propulsion system, Reaction Control System (RCS) for both the Upper Stage and the First Stage. MSFC's Spacecraft and Vehicle Department, Structural and Analysis Design Division is developing a set of predicted mass of these elements. This paper details the methodology, criterion and tools used for the preliminary mass predictions of the upper stage structural assembly components. In general, weight of the cylindrical barrel sections are estimated using the commercial code Hypersizer, whereas, weight of the domes are developed using classical solutions. HyperSizer is software that performs automated structural analysis and sizing optimization based on aerospace methods for strength, stability, and stiffness. Analysis methods range from closed form, traditional hand calculations repeated every day in industry to more advanced panel buckling algorithms. Margin-of-safety reporting for every potential failure provides the engineer with a powerful insight into the structural problem. Optimization capabilities include finding minimum weight panel or beam concepts, material selections, cross sectional dimensions, thicknesses, and lay-ups from a library of 40 different stiffened and sandwich designs and a database of composite, metallic, honeycomb, and foam materials. Multiple different concepts (orthogrid, isogrid, and skin stiffener) were run for multiple loading combinations of ascent design load with and with out tank pressure as well as proof pressure condition. Subsequently, selected optimized concept obtained from Hypersizer runs was translated into a computer aid design (CAD) model to account for the wall thickness tolerance, weld land etc for developing the most probable weight of the components. The flow diram summarizes the analysis steps used in developing these predicted mass.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

A close-up view of the Apollo 11 command service module ready to be mated with the spacecraft LEM adapter of the third stage. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

50 Years of Testing

NASA Image and Video Library

2016-04-23

A 15-second test of a Saturn V rocket stage on the A-2 Test Stand at Stennis Space Center ushered in the Space Age for south Mississippi. Fifty years later, Stennis has grown into the nation’s largest rocket engine test site, continuing to test rocket engines and stages that power the nation’s space program.

77 FR 26943 - Airworthiness Directives; Pratt & Whitney Canada Turboprop Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-08

...) first stage sun gears and planet gears installed in the reduction gearbox. This AD was prompted by failures of certain first stage sun gears manufactured by TAATI. We are issuing this AD to prevent failure of the sun gear and planet gears which will result in an engine in-flight shut down, possible...

76 FR 41430 - Airworthiness Directives; Pratt & Whitney Division (PW) PW4000 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2011-07-14

... removing certain part number (P/N) high-pressure turbine (HPT) stage 1 and HPT stage 2 airseals and HPT... and its high-pressure turbine (HPT). Pratt and Whitney's updated analysis indicated that the current...) Applicability This AD applies to the following Pratt & Whitney Division (PW) turbofan engines, with high...

Apollo 11 Command Service Module

NASA Technical Reports Server (NTRS)

1969-01-01

A close-up view of the Apollo 11 command service module ready to be mated with the spacecraft LEM adapter of the third stage. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Two-stage commercial evaluation of engineering systems production projects for high-rise buildings

NASA Astrophysics Data System (ADS)

Bril, Aleksander; Kalinina, Olga; Levina, Anastasia

2018-03-01

The paper is devoted to the current and debatable problem of methodology of choosing the effective innovative enterprises for venture financing. A two-stage system of commercial innovation evaluation based on the UNIDO methodology is proposed. Engineering systems account for 25 to 40% of the cost of high-rise residential buildings. This proportion increases with the use of new construction technologies. Analysis of the construction market in Russia showed that the production of internal engineering systems elements based on innovative technologies has a growth trend. The production of simple elements is organized in small enterprises on the basis of new technologies. The most attractive for development is the use of venture financing of small innovative business. To improve the efficiency of these operations, the paper proposes a methodology for a two-stage evaluation of small business development projects. A two-stage system of commercial evaluation of innovative projects allows creating an information base for informed and coordinated decision-making on venture financing of enterprises that produce engineering systems elements for the construction business.

SSME Main Combustion Chamber (MCC) hot oil dewaxing

NASA Technical Reports Server (NTRS)

Akpati, Anthony U.

1995-01-01

In an attempt to comply with the changing environmental regulations, a process was developed for the replacement of perchloroethylene in the dewaxing of the Space Shuttle Main Engine (SSME) Main Combustion Chamber (MCC) and other associated hardware filled with the Rigidax (R) casting compound. Rigidax (R) is a hard blue-dyed, calcium carbonate filled thermoplastic casting compound (melting point 77 C) that is melted and poured into hardware cavities to prevent contamination during material removal processes, i.e. machining, grinding, drilling, and deburring. Additionally, it serves as a maskant for designated areas during electroforming processes. Laboratory testing was conducted to evaluate seven alternate fluids for the replacement of perchloroethylene in the dewaxing process. Based upon successful laboratory results, a mineral oil was selected for testing on actual hardware. The final process developed involves simultaneous immersion and flushing of the MCC channels using a distinct eight stage process. A nonvolatile hydrocarbon analysis of a solvent flush sample is performed to determine the hardware cleanliness for comparison to the previous perchloroethylene dewaxing process.

SSME Main Combustion Chamber (MCC) 'Hot Oil' Dewaxing

NASA Technical Reports Server (NTRS)

Akpati, Anthony U.

1994-01-01

In an attempt to comply with the changing environmental regulations, a process was developed for the replacement of perchloroethylene in the dewaxing of the Space Shuttle Main Engine (SSME) Main Combustion Chamber (MCC) and other associated hardware filled with the Rigidax(registered mark) casting compound. Rigidax(registered mark) is a hard blue-dyed, calcium carbonate filled thermoplastic casting compound (melting point 77 C) that is melted and poured into hardware cavities to prevent contamination during material removal processes, i.e. machining, grinding, drilling, and deburring. Additionally, it serves as a maskant for designated areas during electroforming processes. Laboratory testing was conducted to evaluate seven alternate fluids for the replacement of perchloroethylene in the dewaxing process. Based upon successful laboratory results, a mineral oil was selected for testing on actual hardware. The final process developed involves simultaneous immersion and flushing of the MCC channels using a distinct eight stage process. A nonvolatile hydrocarbon analysis of a solvent flush sample is performed to determine the hardware cleanliness for comparison to the previous perchloroethylene dewaxing process.

Planning for Plume Diagnostics for Ground Testing of J-2X Engines at the SSC

NASA Technical Reports Server (NTRS)

SaintCyr, William W.; Tejwani, Gopal D.; McVay, Gregory P.; Langford, Lester A.; SaintCyr, William W.

2010-01-01

John C. Stennis Space Center (SSC) is the premier test facility for liquid rocket engine development and certification for the National Aeronautics and Space Administration (NASA). Therefore, it is no surprise that the SSC will play the most prominent role in the engine development testing and certification for the J-2X engine. The Pratt & Whitney Rocketdyne J-2X engine has been selected by the Constellation Program to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage in NASA s strategy of risk mitigation for hardware development by building on the Apollo program and other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with first demonstration flight in 2010 and human test flights in 2012. Accordingly, J-2X engine design, development, test, and evaluation is to build upon heritage hardware and apply valuable experience gained from past development and testing efforts. In order to leverage SSC s successful and innovative expertise in the plume diagnostics for the space shuttle main engine (SSME) health monitoring,1-10 this paper will present a blueprint for plume diagnostics for various proposed ground testing activities for J-2X at SSC. Complete description of the SSC s test facilities, supporting infrastructure, and test facilities is available in Ref. 11. The A-1 Test Stand is currently being prepared for testing the J-2X engine at sea level conditions. The A-2 Test Stand is currently being used for testing the SSME and may also be used for testing the J-2X engine at sea level conditions in the future. Very recently, ground-breaking ceremony for the new A-3 rocket engine test stand took place at SSC on August 23, 2007. A-3 is the first large - scale test stand to be built at the SSC since the A and B stands were constructed in the 1960s. The A-3 Test Stand will be used for testing J-2X engines under vacuum conditions simulating high altitude operation at approximately 30,480 m (100,000 ft). To achieve the simulated altitude environment, chemical steam generators using isopropyl alcohol, LOX, and RELEASED - Printed documents may be obsolete; validate prior to use. water would run for the duration of the test and would generate approximately 2096 Kg/s of steam to reduce pressure in the test cell and downstream of the engine. The testing at the A-3 Test Stand is projected to begin in late 2010, meanwhile the J-2X component testing on A-1 is scheduled to begin later this year.

NASA Crew and Cargo Launch Vehicle Development Approach Builds on Lessons from Past and Present Missions

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.

2006-01-01

The United States (US) Vision for Space Exploration, announced in January 2004, outlines the National Aeronautics and Space Administration's (NASA) strategic goals and objectives, including retiring the Space Shuttle and replacing it with new space transportation systems for missions to the Moon, Mars, and beyond. The Crew Exploration Vehicle (CEV) that the new human-rated Crew Launch Vehicle (CLV) lofts into space early next decade will initially ferry astronauts to the International Space Station (ISS) Toward the end of the next decade, a heavy-lift Cargo Launch Vehicle (CaLV) will deliver the Earth Departure Stage (EDS) carrying the Lunar Surface Access Module (LSAM) to low-Earth orbit (LEO), where it will rendezvous with the CEV launched on the CLV and return astronauts to the Moon for the first time in over 30 years. This paper outlines how NASA is building these new space transportation systems on a foundation of legacy technical and management knowledge, using extensive experience gained from past and ongoing launch vehicle programs to maximize its design and development approach, with the objective of reducing total life cycle costs through operational efficiencies such as hardware commonality. For example, the CLV in-line configuration is composed of a 5-segment Reusable Solid Rocket Booster (RSRB), which is an upgrade of the current Space Shuttle 4- segment RSRB, and a new upper stage powered by the liquid oxygen/liquid hydrogen (LOX/LH2) J-2X engine, which is an evolution of the J-2 engine that powered the Apollo Program s Saturn V second and third stages in the 1960s and 1970s. The CaLV configuration consists of a propulsion system composed of two 5-segment RSRBs and a 33- foot core stage that will provide the LOX/LED needed for five commercially available RS-68 main engines. The J-2X also will power the EDS. The Exploration Launch Projects, managed by the Exploration Launch Office located at NASA's Marshall Space Flight Center, is leading the design, development, testing, and operations planning for these new space transportation systems. Utilizing a foundation of heritage hardware and management lessons learned mitigates both technical and programmatic risk. Project engineers and managers work closely with the Space Shuttle Program to transition hardware, infrastructure, and workforce assets to the new launch systems, leveraging a wealth of knowledge from Shuffle operations. In addition, NASA and its industry partners have tapped into valuable Apollo databases and are applying corporate wisdom conveyed firsthand by Apollo-era veterans of America s original Moon missions. Learning from its successes and failures, NASA employs rigorous systems engineering and systems management processes and principles in a disciplined, integrated fashion to further improve the probability of mission success.

SLS Engine Section Test Article Moved for Stacking at Michoud

NASA Image and Video Library

2017-04-25

Stacking is underway for the Space Launch System core stage engine section structural qualification test article at NASA's Michoud Assembly Facility in New Orleans. The rocket's engine section is the bottom of the core stage and houses the four RS-25 engines. The engine section test article was moved to Michoud's Cell A in Building 110 for vertical stacking with hardware that simulates the rocket's liquid hydrogen tank, which is the fuel tank that joins to the engine section. Once stacked, the entire test article will load onto the barge Pegasus and ship to NASA's Marshall Space Flight Center in Huntsville, Alabama. There, it will be subjected to millions of pounds of force during testing to ensure the hardware can withstand the incredible stresses of launch.

Tailoring Systems Engineering Processes in a Conceptual Design Environment: A Case Study at NASA Marshall Spaceflight Center's ACO

NASA Technical Reports Server (NTRS)

Mulqueen, John; Maples, C. Dauphne; Fabisinski, Leo, III

2012-01-01

This paper provides an overview of Systems Engineering as it is applied in a conceptual design space systems department at the National Aeronautics and Space Administration (NASA) Marshall Spaceflight Center (MSFC) Advanced Concepts Office (ACO). Engineering work performed in the NASA MFSC's ACO is targeted toward the Exploratory Research and Concepts Development life cycle stages, as defined in the International Council on Systems Engineering (INCOSE) System Engineering Handbook. This paper addresses three ACO Systems Engineering tools that correspond to three INCOSE Technical Processes: Stakeholder Requirements Definition, Requirements Analysis, and Integration, as well as one Project Process Risk Management. These processes are used to facilitate, streamline, and manage systems engineering processes tailored for the earliest two life cycle stages, which is the environment in which ACO engineers work. The role of systems engineers and systems engineering as performed in ACO is explored in this paper. The need for tailoring Systems Engineering processes, tools, and products in the ever-changing engineering services ACO provides to its customers is addressed.

Army Research Office and Air Force Office of Scientific Research Contractors’ Meeting in Chemical Propulsion Held in Boulder, Colorado on June 11-13, 2007

DTIC Science & Technology

2007-06-01

single-cylinder diesel engine showed that at high EGR rates, the autoignition mechanism changes from the typical single stage to a two-stage process...2002). 3. Edwards, T., Zabarnick, S. "Supercritical Fuel Deposition Mechanisms ," Industrial and Engineering Chemistry Research 32: 3117-3122 (1993...and the Air Force Office of Scientific Research. 20070910360 15. SUBJECT TERMS Flames, Propulsion, Gas Turbines, Diesel Engines , Scramjets, Pulse

Saturn V First Stage S-1C LOX Fuel Tanks

NASA Technical Reports Server (NTRS)

1960-01-01

This photograph shows the Saturn V assembled LOX (Liquid Oxygen) and fuel tanks ready for transport from the Manufacturing Engineering Laboratory at Marshall Space Flight Center in Huntsville, Alabama. The tanks were then shipped to the launch site at Kennedy Space Center for a flight. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Cellular membrane enrichment of self-assembling D-peptides for cell surface engineering.

PubMed

Wang, Huaimin; Wang, Youzhi; Han, Aitian; Cai, Yanbin; Xiao, Nannan; Wang, Ling; Ding, Dan; Yang, Zhimou

2014-06-25

We occasionally found that several self-assembling peptides containing D-amino acids would be preferentially enriched in cellular membranes at self-assembled stages while distributed evenly in the cytoplasma of cells at unassembled stages. Self-assembling peptides containing only Lamino acids distributed evenly in cytoplasma of cells at both self-assembled and unassembled stages. The self-assembling peptides containing D-amino acids could therefore be applied for engineering cell surface with peptides. More importantly, by integrating a protein binding peptide (a PDZ domain binding hexapeptide of WRESAI) with the self-assembling peptide containing D-amino acids, protein could also be introduced to the cell surface. This study not only provided a novel approach to engineer cell surface, but also highlighted the unusual properties and potential applications of self-assembling peptides containing D-amino acids in regenerative medicine, drug delivery, and tissue engineering.

Lessons Learned from Ares I Upper Stage Structures and Thermal Design

NASA Technical Reports Server (NTRS)

Ahmed, Rafiq

2012-01-01

The Ares 1 Upper Stage was part of the vehicle intended to succeed the Space Shuttle as the United States manned spaceflight vehicle. Although the Upper Stage project was cancelled, there were many lessons learned that are applicable to future vehicle design. Lessons learned that are briefly detailed in this Technical Memorandum are for specific technical areas such as tank design, common bulkhead design, thrust oscillation, control of flight and slosh loads, purge and hazardous gas system. In addition, lessons learned from a systems engineering and vehicle integration perspective are also included, such as computer aided design and engineering, scheduling, and data management. The need for detailed systems engineering in the early stages of a project is emphasized throughout this report. The intent is that future projects will be able to apply these lessons learned to keep costs down, schedules brief, and deliver products that perform to the expectations of their customers.

NASA Conducts First RS-25 Rocket Engine Test of 2015

NASA Image and Video Library

2015-01-09

From the Press Release: The new year is off to a hot start for NASA's Space Launch System (SLS). The engine that will drive America's next great rocket to deep space blazed through its first successful test Jan. 9 at the agency's Stennis Space Center near Bay St. Louis, Mississippi. The RS-25, formerly the space shuttle main engine, fired up for 500 seconds on the A-1 test stand at Stennis, providing NASA engineers critical data on the engine controller unit and inlet pressure conditions. This is the first hot fire of an RS-25 engine since the end of space shuttle main engine testing in 2009. Four RS-25 engines will power SLS on future missions, including to an asteroid and Mars. "We’ve made modifications to the RS-25 to meet SLS specifications and will analyze and test a variety of conditions during the hot fire series,” said Steve Wofford, manager of the SLS Liquid Engines Office at NASA's Marshall Space Flight Center in Huntsville, Alabama, where the SLS Program is managed. "The engines for SLS will encounter colder liquid oxygen temperatures than shuttle; greater inlet pressure due to the taller core stage liquid oxygen tank and higher vehicle acceleration; and more nozzle heating due to the four-engine configuration and their position in-plane with the SLS booster exhaust nozzles.” The engine controller unit, the "brain" of the engine, allows communication between the vehicle and the engine, relaying commands to the engine and transmitting data back to the vehicle. The controller also provides closed-loop management of the engine by regulating the thrust and fuel mixture ratio while monitoring the engine's health and status. The new controller will use updated hardware and software configured to operate with the new SLS avionics architecture. "This first hot-fire test of the RS-25 engine represents a significant effort on behalf of Stennis Space Center’s A-1 test team," said Ronald Rigney, RS-25 project manager at Stennis. "Our technicians and engineers have been working diligently to design, modify and activate an extremely complex and capable facility in support of RS-25 engine testing." Testing will resume in April after upgrades are completed on the high pressure industrial water system, which provides cool water for the test facility during a hot fire test. Eight tests, totaling 3,500 seconds, are planned for the current development engine. Another development engine later will undergo 10 tests, totaling 4,500 seconds. The second test series includes the first test of new flight controllers, known as green running. The first flight test of the SLS will feature a configuration for a 70-metric-ton (77-ton) lift capacity and carry an uncrewed Orion spacecraft beyond low-Earth orbit to test the performance of the integrated system. As the SLS is upgraded, it will provide an unprecedented lift capability of 130 metric tons (143 tons) to enable missions even farther into our solar system.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

The fuel tank assembly of the Saturn V S-IC (first) stage supported with the aid of a C frame on the transporter was readied to be transported to the Marshall Space Flight Center, building 4705. The fuel tank carried kerosene (RP-1) as its fuel. The S-IC stage utilized five F-1 engines that used kerosene and liquid oxygen as propellant and each engine provided 1,500,000 pounds of thrust. This stage lifted the entire vehicle and Apollo spacecraft from the launch pad.

Saturn Apollo Program

NASA Image and Video Library

1964-11-01

The Saturn I S-IV stage (second stage) assembly for the SA-9 mission underwent the weight and balance test in the hanger building at Cape Canaveral. The S-IV stage had six RL-10 engines which used liquid hydrogen and liquid oxygen as its propellants arranged in a circle. Each RL-10 engine produced a thrust of 15,000 pounds, a total combined thrust of 90,000 pounds. The SA-9 mission was the first Saturn with operational payload Pegasus I, meteoroid detection satellite, and launched on February 16, 1965.

Saturn Apollo Program

NASA Image and Video Library

1967-11-01

This image depicts a Boeing worker installing an F-1 engine on the Saturn V S-IC flight stage at the Michoud Assembly Facility (MAF). The Saturn IB and Saturn V first stages were manufactured at the MAF, located 24 kilometers (approximately 15 miles) east of downtown New Orleans, Louisiana. The prime contractors, Chrysler and Boeing, jointly occupied the MAF. The basic manufacturing building boasted 43 acres under one roof. By 1964, NASA added a separate engineering and office building, vertical assembly building, and test stage building.

Brush seal numerical simulation: Concepts and advances

NASA Technical Reports Server (NTRS)

Braun, M. J.; Kudriavtsev, V. V.

1994-01-01

The development of the brush seal is considered to be most promising among the advanced type seals that are presently in use in the high speed turbomachinery. The brush is usually mounted on the stationary portions of the engine and has direct contact with the rotating element, in the process of limiting the 'unwanted' leakage flows between stages, or various engine cavities. This type of sealing technology is providing high (in comparison with conventional seals) pressure drops due mainly to the high packing density (around 100 bristles/sq mm), and brush compliance with the rotor motions. In the design of modern aerospace turbomachinery leakage flows between the stages must be minimal, thus contributing to the higher efficiency of the engine. Use of the brush seal instead of the labyrinth seal reduces the leakage flow by one order of magnitude. Brush seals also have been found to enhance dynamic performance, cost less, and are lighter than labyrinth seals. Even though industrial brush seals have been successfully developed through extensive experimentation, there is no comprehensive numerical methodology for the design or prediction of their performance. The existing analytical/numerical approaches are based on bulk flow models and do not allow the investigation of the effects of brush morphology (bristle arrangement), or brushes arrangement (number of brushes, spacing between them), on the pressure drops and flow leakage. An increase in the brush seal efficiency is clearly a complex problem that is closely related to the brush geometry and arrangement, and can be solved most likely only by means of a numerically distributed model.

Attitudes towards chemistry among engineering students

NASA Astrophysics Data System (ADS)

Olivo Delgado, Carlos J.

The attitudes towards chemistry of the engineering students enrolled in an introductory course at the Polytechnic University of Puerto Rico were explored (n = 115). A mixed methodology was used in an exploratory-oriented research approach. The first stage consisted of the administration of a Likert scale attitudinal survey which was validated during the study's design process. The survey allowed collecting information about the participant's attitudes towards their personal opinion, their perspectives about peer's or relatives' opinion, relevant aspects of the discipline, and difficulty-increasing factors in the chemistry course. The scale internal reliability was measured in a pilot study with a convenience simple, obtaining an acceptable coefficient (Cronbach alpha = 0.731). Survey results evidenced a mainly neutral attitude towards the chemistry course, not highly negative or highly positive, in contrast with other studies in this field. On the other hand, the normality hypothesis was tested for the scores obtained by participants in the survey. Although the pilot study sample had an approximately normal distribution, the scores in obtained by the participants in the survey failed the normality test criteria. The second stage of the study was accomplished using a case study. Among the survey participants, some students were invited to in-depth interviews to elucidate the reasons why they have certain attitudes towards chemistry. Study time, instructor, motivation, term of study, and course schedule are the factors that interviewees agreed as contributors to success or failure in the chemistry course. Interview's participants emphasized that study time is determinant to pass the class. This methodological approach, quantitative followed by qualitative, was useful in describing the attitudes towards chemistry among university students of engineering.

Design of Smart Educational Robot as a Tool For Teaching Media Based on Contextual Teaching and Learning to Improve the Skill of Electrical Engineering Student

NASA Astrophysics Data System (ADS)

Zuhrie, M. S.; Basuki, I.; Asto, B. I. G. P.; Anifah, L.

2018-04-01

The development of robotics in Indonesia has been very encouraging. The barometer is the success of the Indonesian Robot Contest. The focus of research is a teaching module manufacturing, planning mechanical design, control system through microprocessor technology and maneuverability of the robot. Contextual Teaching and Learning (CTL) strategy is the concept of learning where the teacher brings the real world into the classroom and encourage students to make connections between knowledge possessed by its application in everyday life. This research the development model used is the 4-D model. This Model consists of four stages: Define Stage, Design Stage, Develop Stage, and Disseminate Stage. This research was conducted by applying the research design development with the aim to produce a tool of learning in the form of smart educational robot modules and kit based on Contextual Teaching and Learning at the Department of Electrical Engineering to improve the skills of the Electrical Engineering student. Socialization questionnaires showed that levels of the student majoring in electrical engineering competencies image currently only limited to conventional machines. The average assessment is 3.34 validator included in either category. Modules developed can give hope to the future are able to produce Intelligent Robot Tool for Teaching.

Developing the STS sound pollution unit for enhancing students' applying knowledge among science technology engineering and mathematics

NASA Astrophysics Data System (ADS)

Jumpatong, Sutthaya; Yuenyong, Chokchai

2018-01-01

STEM education suggested that students should be enhanced to learn science with integration between Science, Technology, Engineering and Mathematics. To help Thai students make sense of relationship between Science, Technology, Engineering and Mathematics, this paper presents learning activities of STS Sound Pollution. The developing of STS Sound Pollution is a part of research that aimed to enhance students' perception of the relationship between Science Technology Engineering and Mathematics. This paper will discuss how to develop Sound Pollution through STS approach in framework of Yuenyong (2006) where learning activities were provided based on 5 stages. These included (1) identification of social issues, (2) identification of potential solutions, (3) need for knowledge, (4) decisionmaking, and (5) socialization stage. The learning activities could be highlighted as following. First stage, we use video clip of `Problem of people about Sound Pollution'. Second stage, students will need to identification of potential solutions by design Home/Factory without noisy. The need of scientific and other knowledge will be proposed for various alternative solutions. Third stage, students will gain their scientific knowledge through laboratory and demonstration of sound wave. Fourth stage, students have to make decision for the best solution of designing safety Home/Factory based on their scientific knowledge and others (e.g. mathematics, economics, art, value, and so on). Finally, students will present and share their Design Safety Home/Factory in society (e.g. social media or exhibition) in order to validate their ideas and redesigning. The paper, then, will discuss how those activities would allow students' applying knowledge of science technology engineering, mathematics and others (art, culture and value) for their possible solution of the STS issues.

NASA's Space Launch System: Progress Report

NASA Technical Reports Server (NTRS)

Cook, Jerry; Lyles, Garry

2017-01-01

NASA and its commercial industry team achieved significant progress in 2016 in manufacturing and testing of the Block 1 vehicle for the first launch of the Space Launch System (SLS). Test and flight article hardware for the liquid hydrogen fuel tank as well as the engine section for the core stage were completed at Michoud Assembly Facility (MAF) in New Orleans. Test stands neared completion at Marshall Space Flight Center for the propellant tanks, engine section, intertank and payload section. Stennis Space Center completed major structural renovations on the B2 test stand, where the core stage "green run" test program will be conducted. The SLS team completed a hotfire test series at Stennis to successfully demonstrate the ability of the RS-25 engine to operate under SLS environments and performance conditions. The team also test fired the second qualification five-segment solid rocket motor and cast the first six motor segments for the first SLS mission. The Interim Cryogenic Propulsion Stage (ICPS) test article was delivered to Marshall for structural tests, and work is nearly finished on the flight stage. Flight software testing completed at Marshall included power quality and command and data handling. In 2017, that work continues. SLS completed Preliminary Design Review (PDR) on the Exploration Upper Stage (EUS), a powerful, human-rated spacecraft that will propel explorers to cis-lunar space. In 2017, hardware will continue to be integrated at MAF for core stage structural test articles and the first two operational flights. RS-25 hotfire testing will continue to explore engine performance, as well as test flight-like software and four new Engine Controller Units (ECUs) for the first mission. Production of development components for a more affordable RS-25 design is underway. Core stage structural test articles have begun arriving at Marshall. While engineering challenges typical of a new development are possible, SLS is working toward launch readiness in late 2018. This paper will discuss these and other technical and programmatic successes and challenges over the past year and provide a preview of work ahead before first flight

Production, purification and utilization of biogas as fuel for internal combustion engine

NASA Astrophysics Data System (ADS)

Hernandez, Noel M.; Villanueva, Eliseo P.

2018-03-01

This study attempts to modify a 4-cylinder gasoline engine to run with a purified compressed biogas as substitute for fossil fuels. Water scrubbing method was used as the easiest purification technique to remove CO2 and iron filing for H2S. The pressurized raw biogas was fed in a low cost made portable floating type gas holder with volume capacity of 0.74 m3. The purified biogas was compressed using a reciprocating compressor through a two stage series of enrichment and moisture removal process using activated alumina into the steel cylinder to improve the quality of the methane content. The enriched biogas was filled in the LPG tank for 20 minutes at 10 bars at an average of 73.67% CH4 with no traces of H2S as storage for engine utilization. The modification involved the installation and mounting of LPG conversion kit. A comparative analysis of the performance and combustion characteristics of the engine was evaluated separately with gasoline and purified compressed biogas using electro-dynamometer as variable loads. The findings show that power output deterioration in compressed biogas was mainly due to high percentage of CO2 and other gases impurities. It also shows that because of the calorific value of biogas, the thermal efficiency is lesser than that of gasoline. It implies that the overall engine performance can be improved by removing undesirable gases in the mixture.

NACA Researcher Measures Ice on a Turbojet Engine Inlet

NASA Image and Video Library

1948-11-21

The National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory conducted an extensive icing research program in the late 1940s that included studies in the Icing Research Tunnel and using specially modified aircraft. One facet of this program was the investigation of the effects of icing on turbojets. Although jet engines allowed aircraft to pass through inclement weather at high rates of speed, ice accumulation was still a concern. The NACA’s B-24M Liberator was initially reconfigured with a General Electric I-16 engine installed in the aircraft’s waist compartment with an air scoop and spray nozzles to produce the artificial icing conditions. The centrifugal engine appeared nearly impervious to the effects of icing. Axial-flow jet engines, however, were much more susceptible to icing damage. The inlet guide vanes were particularly vulnerable, but the cowling’s leading edge, the main bearing supports, and accessory housing could also ice up. If pieces of ice reached the engine’s internal components, the compressor blades could be damaged. To study this phenomenon, a Westinghouse 24C turbojet, seen in this photograph, was installed under the B-24M’s right wing. In January 1948 flight tests of the 24C in icing conditions began. Despite ice buildup into the second stage of the compressor, the engine was able to operate at takeoff speeds. Researchers found the ice on the inlet vanes resulted in half of the engine’s decreased performance.

Key fuel properties and engine performances of diesel-ethanol blends, using tetrahydrofuran as surfactant additive

NASA Astrophysics Data System (ADS)

Molea, A.; Visuian, P.; Barabás, I.; Suciu, R. C.; Burnete, N. V.

2017-10-01

In this paper there were presented researches related to preparation and characterization of physicochemical properties of diesel-ethanol blends stabilized with tetrahydrofuran as surfactant, in order to be used as fuels in compression ignition engines. The main spray characteristics and engine performances of these blends were evaluated by using AVL Fire software. In the first stage of the studies, commercial diesel was mixed with ethanol, in different concentrations (between 2% and 15% v/v), followed by the addition of tetrahydrofuran (THF) until the blends were miscible, i.e. the blends were stabilized. The experiments were done at room temperature (22 °C). The obtained blends were characterized in order to determine the chemical composition and physicochemical properties, i.e. density, kinematic viscosity, surface tension. UV-Vis spectroscopy was utilized in order to determine a semi-quantitative evaluation regarding the chemical composition of the prepared blends and chemical interaction between diesel, ethanol and THF. Based on the determined properties, the fuel spray characteristics, engine performances and emission characteristics were evaluated by simulation using the AVL Fire software. The obtained results regarding physicochemical properties of blends were compared with diesel. Some improvements were observed when operating with the prepared blends compared to diesel with respect to engine performances and emission characteristics. Based on physicochemical evaluation and computer simulation, it was demonstrated that diesel-ethanol-tetrahydrofuran blends can be used as alternative fuel in compression ignition engines.

Development of advanced high temperature in-cylinder components and tribological systems for low heat rejection diesel engines, phase 1

NASA Astrophysics Data System (ADS)

Kroeger, C. A.; Larson, H. J.

1992-03-01

Analysis and concept design work completed in Phase 1 have identified a low heat rejection engine configuration with the potential to meet the Heavy Duty Transport Technology program specific fuel consumption goal of 152 g/kW-hr. The proposed engine configuration incorporates low heat rejection, in-cylinder components designed for operation at 24 MPa peak cylinder pressure. Water cooling is eliminated by selective oil cooling of the components. A high temperature lubricant will be required due to increased in-cylinder operating temperatures. A two-stage turbocharger air system with intercooling and aftercooling was selected to meet engine boost and BMEP requirements. A turbocompound turbine stage is incorporated for exhaust energy recovery. The concept engine cost was estimated to be 43 percent higher compared to a Caterpillar 3176 engine. The higher initial engine cost is predicted to be offset by reduced operating costs due the lower fuel consumption.

Development of advanced high temperature in-cylinder components and tribological systems for low heat rejection diesel engines, phase 1

NASA Technical Reports Server (NTRS)

Kroeger, C. A.; Larson, H. J.

1992-01-01

Analysis and concept design work completed in Phase 1 have identified a low heat rejection engine configuration with the potential to meet the Heavy Duty Transport Technology program specific fuel consumption goal of 152 g/kW-hr. The proposed engine configuration incorporates low heat rejection, in-cylinder components designed for operation at 24 MPa peak cylinder pressure. Water cooling is eliminated by selective oil cooling of the components. A high temperature lubricant will be required due to increased in-cylinder operating temperatures. A two-stage turbocharger air system with intercooling and aftercooling was selected to meet engine boost and BMEP requirements. A turbocompound turbine stage is incorporated for exhaust energy recovery. The concept engine cost was estimated to be 43 percent higher compared to a Caterpillar 3176 engine. The higher initial engine cost is predicted to be offset by reduced operating costs due the lower fuel consumption.

Saturn Apollo Program

NASA Image and Video Library

1965-03-01

The hydrogen-powered second stage is being lowered into place during the final phase of fabrication of the Saturn V moon rocket at North American's Seal Beach, California facility. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

Workers at the Michoud Assembly Facility near New Orleans, Louisiana install the H-1 engines into the S-IB stage, the Saturn IB launch vehicle's first stage. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at MAF, the 90,000-pound booster utilized eight H-1 engines to produce a combined thrust of 1,600,000 pounds.

Performance and control study of a low-pressure-ratio turbojet engine for a drone aircraft

NASA Technical Reports Server (NTRS)

Seldner, K.; Geyser, L. C.; Gold, H.; Walker, D.; Burgner, G.

1972-01-01

The results of analog and digital computer studies of a low-pressure-ratio turbojet engine system for use in a drone vehicle are presented. The turbojet engine consists of a four-stage axial compressor, single-stage turbine, and a fixed area exhaust nozzle. Three simplified fuel schedules and a generalized parameter fuel control for the engine system are presented and evaluated. The evaluation is based on the performance of each schedule or control during engine acceleration from a windmill start at Mach 0.8 and 6100 meters to 100 percent corrected speed. It was found that, because of the higher acceleration margin permitted by the control, the generalized parameter control exhibited the best dynamic performance.

Wernher von Braun

NASA Image and Video Library

2004-04-15

A pioneer of America's space program, Dr. von Braun stands by the five F-1 engines of the Saturn V launch vehicle. This Saturn V vehicle is an actual test vehicle which has been displayed at the U.S. Space Rocket Center in Huntsville, Alabama. Designed and developed by Rocketdyne under the direction of the Marshall Space Flight Center, a cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage. The engines measured 19-feet tall by 12.5-feet at the nozzle exit and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust. The S-IC stage is the first stage, or booster, of a 364-foot long rocket that ultimately took astronauts to the Moon.

Single shaft automotive gas turbine engine characterization test

NASA Technical Reports Server (NTRS)

Johnson, R. A.

1979-01-01

An automotive gas turbine incorporating a single stage centrifugal compressor and a single stage radial inflow turbine is described. Among the engine's features is the use of wide range variable geometry at the inlet guide vanes, the compressor diffuser vanes, and the turbine inlet vanes to achieve improved part load fuel economy. The engine was tested to determine its performance in both the variable geometry and equivalent fixed geometry modes. Testing was conducted without the originally designed recuperator. Test results were compared with the predicted performance of the nonrecuperative engine based on existing component rig test maps. Agreement between test results and the computer model was achieved.

Progress on the J-2X Upper Stage Engine for the Ares I Crew Launch Vehicle and the Ares V Cargo Launch Vehicle

NASA Technical Reports Server (NTRS)

Byrd, Thomas D.; Kynard, Michael .

2007-01-01

NASA's Vision for Exploration requires a safe, reliable, affordable upper stage engine to power the Ares I Crew Launch Vehicle (CLV) and the Ares V Cargo Launch Vehicle. The J-2X engine is being developed for that purpose, epitomizing NASA's philosophy of employing legacy knowledge, heritage hardware, and commonality to carry the next generation of explorers into low-Earth orbit and out into the solar system This presentation gives top-level details on accomplishments to date and discusses forward work necessary to bring the J-2X engine to the launch pad.

An engine awaits processing in the new engine shop at KSC

NASA Technical Reports Server (NTRS)

1998-01-01

In the Space Shuttle Main Engine Processing Facility (SSMEPF), a new Block 2A engine sits on the workstand as technicians process it. The engine is scheduled to fly on the Space Shuttle Endeavour during the STS-88 mission in December 1998. The SSMEPF officially opened on July 6, replacing the Shuttle Main Engine Shop.

Fast Track NTR Systems Assessment for NASA's First Lunar Outpost Scenario

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Alexander, Stephen W.

1994-01-01

Integrated systems and mission study results are presented which quantify the rationale and benefits for developing and using nuclear thermal rocket (NTR) technology for returning humans to the moon in the early 2000's. At present, the Exploration Program Office (ExPO) is considering chemical propulsion for its 'First Lunar Outpost' (FLO) mission, and NTR propulsion for the more demanding Mars missions to follow. The use of an NTR-based lunar transfer stage, capable of evolving to Mars mission applications, could result in an accelerated schedule, reduced cost approach to moon/Mars exploration. Lunar mission applications would also provide valuable operational experience and serve as a 'proving ground' for NTR engine and stage technologies. In terms of performance benefits, studies indicate that an expendable NTR stage powered by two 50 klbf engines can deliver approximately 96 metric tons (t) to trans-lunar injection (TLI) conditions for an initial mass in low earth orbit (IMLEO) of approximately 199 t compared to 250 t for a cryogenic chemical TLI stage. The NTR stage liquid hydrogen (LH2) tank has a 10 m diameter, 14.8 m length, and 68 t LH2 capacity. The NTR utilizes a 'graphite' fuel form consisting of coated UC2 particles in a graphite substrate, and has a specific impulse capability of approximately 870 s, and an engine thrust-to-weight ratio of approximately 4.8. The NTR stage and its piloted FLO lander has a total length of approximately 38 m and can be launched by a single Saturn V-derived heavy lift launch vehicle (HLLV) in the 200 to 250 t-class range. The paper summarizes NASA's First Lunar Outpost scenario, describes characteristics for representative engine/stage configurations, and examines the impact on engine selection and vehicle design resulting from a consideration of alternative NTR fuel forms and lunar mission profiles.

The Engineering Development Array: A Low Frequency Radio Telescope Utilising SKA Precursor Technology

NASA Astrophysics Data System (ADS)

Wayth, Randall; Sokolowski, Marcin; Booler, Tom; Crosse, Brian; Emrich, David; Grootjans, Robert; Hall, Peter J.; Horsley, Luke; Juswardy, Budi; Kenney, David; Steele, Kim; Sutinjo, Adrian; Tingay, Steven J.; Ung, Daniel; Walker, Mia; Williams, Andrew; Beardsley, A.; Franzen, T. M. O.; Johnston-Hollitt, M.; Kaplan, D. L.; Morales, M. F.; Pallot, D.; Trott, C. M.; Wu, C.

2017-08-01

We describe the design and performance of the Engineering Development Array, which is a low-frequency radio telescope comprising 256 dual-polarisation dipole antennas working as a phased array. The Engineering Development Array was conceived of, developed, and deployed in just 18 months via re-use of Square Kilometre Array precursor technology and expertise, specifically from the Murchison Widefield Array radio telescope. Using drift scans and a model for the sky brightness temperature at low frequencies, we have derived the Engineering Development Array's receiver temperature as a function of frequency. The Engineering Development Array is shown to be sky-noise limited over most of the frequency range measured between 60 and 240 MHz. By using the Engineering Development Array in interferometric mode with the Murchison Widefield Array, we used calibrated visibilities to measure the absolute sensitivity of the array. The measured array sensitivity matches very well with a model based on the array layout and measured receiver temperature. The results demonstrate the practicality and feasibility of using Murchison Widefield Array-style precursor technology for Square Kilometre Array-scale stations. The modular architecture of the Engineering Development Array allows upgrades to the array to be rolled out in a staged approach. Future improvements to the Engineering Development Array include replacing the second stage beamformer with a fully digital system, and to transition to using RF-over-fibre for the signal output from first stage beamformers.

Status of the Combined Cycle Engine Rig

NASA Technical Reports Server (NTRS)

Saunders, Dave; Slater, John; Dippold, Vance

2009-01-01

Status for the past year is provided of the turbine-based Combined-Cycle Engine (CCE) Rig for the hypersonic project. As part of the first stage propulsion of a two-stage-to-orbit vehicle concept, this engine rig is designed with a common inlet that supplies flow to a turbine engine and a dual-mode ramjet / scramjet engine in an over/under configuration. At Mach 4 the inlet has variable geometry to switch the airflow from the turbine to the ramjet / scramjet engine. This process is known as inlet mode-transition. In addition to investigating inlet aspects of mode transition, the rig will allow testing of turbine and scramjet systems later in the test series. Fully closing the splitter cowl "cocoons" the turbine engine and increases airflow to the scramjet duct. The CCE Rig will be a testbed to investigate integrated propulsion system and controls technology objectives. Four phases of testing are planned to 1) characterize the dual inlet database, 2) collect inlet dynamics using system identification techniques, 3) implement an inlet control to demonstrate mode-transition scenarios and 4) demonstrate integrated inlet/turbine engine operation through mode-transition. Status of the test planning and preparation activities is summarized with background on the inlet design and small-scale testing, analytical CFD predictions and some details of the large-scale hardware. The final stages of fabrication are underway.

Prediction of force coefficients for labyrinth seals

NASA Technical Reports Server (NTRS)

Lee, O. W. K.; Martinez-Sanchez, M.; Czajkowski, E.

1984-01-01

The development of a linear model for the prediction of labyrinth seal forces and on its comparison to available stiffness data is presented. A discussion of the relevance of fluid damping forces and the preliminary stages of a program to obtain data on these forces are examined. Fluid-dynamic forces arising from nonuniform pressure patterns in labyrinth seal glands are known to be potentially destablizing in high power turbomachinery. A well documented case in point is that of the space Shuttle Main Engine turbopumps. Seal forces are also an important factor for the stability of shrouded turbines, acting in that case in conjunction with the effects of blade-tip clearance variations.

Thermal stratification in LH2 tank of cryogenic propulsion stage tested in ISRO facility

NASA Astrophysics Data System (ADS)

Xavier, M.; Raj, R. Edwin; Narayanan, V.

2017-02-01

Liquid oxygen and hydrogen are used as oxidizer and fuel respectively in cryogenic propulsion system. These liquids are stored in foam insulated tanks of cryogenic propulsion system and are pressurized using warm pressurant gas supplied for tank pressure maintenance during cryogenic engine operation. Heat leak to cryogenic propellant tank causes buoyancy driven liquid stratification resulting in formation of warm liquid stratum at liquid free surface. This warm stratum is further heated by the admission of warm pressurant gas for tank pressurization during engine operation. Since stratified layer temperature has direct bearing on the cavitation free operation of turbo pumps integrated in cryogenic engine, it is necessary to model the thermal stratification for predicting stratified layer temperature and mass of stratified liquid in tank at the end of engine operation. These inputs are required for estimating the minimum pressure to be maintained by tank pressurization system. This paper describes configuration of cryogenic stage for ground qualification test, stage hot test sequence, a thermal model and its results for a foam insulated LH2 tank subjected to heat leak and pressurization with hydrogen gas at 200 K during liquid outflow at 38 lps for engine operation. The above model considers buoyancy flow in free convection boundary layer caused by heat flux from tank wall and energy transfer from warm pressurant gas etc. to predict temperature of liquid stratum and mass of stratified liquid in tank at the end of engine operation in stage qualification tests carried out in ISRO facility.

Booster propulsion/vehicle impact study, 2

NASA Technical Reports Server (NTRS)

Johnson, P.; Satterthwaite, S.; Carson, C.; Schnackel, J.

1988-01-01

This is the final report in a study examining the impact of launch vehicles for various boost propulsion design options. These options included: differing boost phase engines using different combinations of fuels and coolants to include RP-1, methane, propane (subcooled and normal boiling point), and hydrogen; variable and high mixture ratio hydrogen engines; translating nozzles on boost phase engines; and cross feeding propellants from the booster to second stage. Vehicles examined included a fully reusable two stage cargo vehicle and a single stage to orbit vehicle. The use of subcooled propane as a fuel generated vehicles with the lowest total vehicle dry mass. Engines with hydrogen cooling generated only slight mass reductions from the reference, all-hydrogen vehicle. Cross feeding propellants generated the most significant mass reductions from the reference two stage vehicle. The use of high mixture ratio or variable mixture ratio hydrogen engines in the boost phase of flight resulted in vehicles with total dry mass 20 percent greater than the reference hydrogen vehicle. Translating nozzles for boost phase engines generated a heavier vehicle. Also examined were the design impacts on the vehicle and ground support subsystems when subcooled propane is used as a fuel. The most significant cost difference between facilities to handle normal boiling point versus subcooled propane is 5 million dollars. Vehicle cost differences were negligible. A significant technical challenge exists for properly conditioning the vehicle propellant on the ground and in flight when subcooled propane is used as fuel.

Testing to Transition the J-2X from Paper to Hardware

NASA Technical Reports Server (NTRS)

Byrd, Tom

2010-01-01

The J-2X Upper Stage Engine (USE) will be the first new human-rated upper stage engine since the Apollo program of the 1960s. It is designed to carry the Ares I and Ares V into orbit and send the Ares V to the Moon as part of NASA's Constellation Program. This paper will provide an overview of progress on the design, testing, and manufacturing of this new engine in 2009 and 2010. The J-2X embodies the program goals of basing the design on proven technology and experience and seeking commonality between the Ares vehicles as a way to minimize risk, shorten development times, and live within current budget constraints. It is based on the proven J-2 engine used on the Saturn IB and Saturn V launch vehicles. The prime contractor for the J-2X is Pratt & Whitney Rocketdyne (PWR), which is under a design, development, test, and engineering (DDT&E) contract covering the period from June 2006 through September 2014. For Ares I, the J-2X will provide engine start at approximately 190,000 feet, operate roughly 500 seconds, and shut down. For Ares V, the J-2X will start at roughly 190,000 feet to place the Earth departure stage (EDS) in orbit, shut down and loiter for up to five days, re-start on command and operate for roughly 300 seconds at its secondary power level to perform trans lunar injection (TLI), followed by final engine shutdown. The J-2X development effort focuses on four key areas: early risk mitigation, design risk mitigation, component and subassembly testing, and engine system testing. Following that plan, the J-2X successfully completed its critical design review (CDR) in 2008, and it has made significant progress in 2009 and 2010 in moving from the drawing board to the machine shop and test stand. Post-CDR manufacturing is well under way, including PWR in-house and vendor hardware. In addition, a wide range of component and sub-component tests have been completed, and more component tests are planned. Testing includes heritage powerpack, turbopump inducer water flow, turbine air flow, turbopump seal testing, main injector and gas generator, injector testing, augmented spark igniter testing, nozzle side loads cold flow testing, nozzle extension film cooling flow testing, control system testing with hardware in the loop, and nozzle extension emissivity coating tests. In parallel with hardware manufacturing, work is progressing on the new A-3 test stand to support full duration altitude testing. The Stennis A-2 test stand is scheduled to be turned over to the Constellation Program in September 2010 to be modified for J-2X testing also. As the structural steel was rising on the A-3 stand, work was under way in the nearby E complex on the chemical steam generator and subscale diffuser concepts to be used to evacuate the A-3 test cell and simulate altitude conditions.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This cutaway illustration shows the Saturn V S-IVB (third) stage with the callouts of its major components. When the S-II (second) stage of the powerful Saturn V rocket burnt out and was separated the remaining units approached orbit around the Earth. Injection into the desired orbit was attaineded as the S-IVB (third stage) was ignited and burnt. The S-IVB stage was powered by a single 200,000-pound thrust J-2 engine and had a re-start capability built in for its J-2 engine. The S-IVB restarted to speed the Apollo spacecraft to escape velocity injecting it and the astronauts into a moon trajectory.

Dynamic Modeling of Starting Aerodynamics and Stage Matching in an Axi-Centrifugal Compressor

NASA Technical Reports Server (NTRS)

Wilkes, Kevin; OBrien, Walter F.; Owen, A. Karl

1996-01-01

A DYNamic Turbine Engine Compressor Code (DYNTECC) has been modified to model speed transients from 0-100% of compressor design speed. The impetus for this enhancement was to investigate stage matching and stalling behavior during a start sequence as compared to rotating stall events above ground idle. The model can simulate speed and throttle excursions simultaneously as well as time varying bleed flow schedules. Results of a start simulation are presented and compared to experimental data obtained from an axi-centrifugal turboshaft engine and companion compressor rig. Stage by stage comparisons reveal the front stages to be operating in or near rotating stall through most of the start sequence. The model matches the starting operating line quite well in the forward stages with deviations appearing in the rearward stages near the start bleed. Overall, the performance of the model is very promising and adds significantly to the dynamic simulation capabilities of DYNTECC.