Advanced engine study for mixed-mode orbit-transfer vehicles

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1978-01-01

Engine design, performance, weight and envelope data were established for three mixed-mode orbit-transfer vehicle engine candidates. Engine concepts evaluated are the tripropellant, dual-expander and plug cluster. Oxygen, RP-1 and hydrogen are the propellants considered for use in these engines. Theoretical performance and propellant properties were established for bipropellant and tripropellant mixes of these propellants. RP-1, hydrogen and oxygen were evaluated as coolants and the maximum attainable chamber pressures were determined for each engine concept within the constraints of the propellant properties and the low cycle thermal fatigue (300 cycles) requirement. The baseline engine design and component operating characteristics are determined at a thrust level of 88,964N (20,000 lbs) and a thrust split of 0.5. The parametric data is generated over ranges of thrust and thrust split of 66.7 to 400kN (15 to 90 klb) and 0.4 to 0.8, respectively.

Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Cataldo, R. L.

2001-01-01

The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

Low heat transfer oxidizer heat exchanger design and analysis

NASA Technical Reports Server (NTRS)

Kanic, P. G.; Kmiec, T. D.; Peckham, R. J.

1987-01-01

The RL10-IIB engine, a derivative of the RLIO, is capable of multi-mode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2 percent of full thrust, and pumped idle (PI), which is 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for log-g deployment. Stable combustion of the RL10-IIB engine at THI and PI thrust levels can be accomplished by providing gaseous oxygen at the propellant injector. Using gaseous hydrogen from the thrust chamber jacket as an energy source, a heat exchanger can be used to vaporize liquid oxygen without creating flow instability. This report summarizes the design and analysis of a United Aircraft Products (UAP) low-rate heat transfer heat exchanger concept for the RL10-IIB rocket engine. The design represents a second iteration of the RL10-IIB heat exchanger investigation program. The design and analysis of the first heat exchanger effort is presented in more detail in NASA CR-174857. Testing of the previous design is detailed in NASA CR-179487.

Some Calculated Research Results of the Working Process Parameters of the Low Thrust Rocket Engine Operating on Gaseous Oxygen-Hydrogen Fuel

NASA Astrophysics Data System (ADS)

Ryzhkov, V.; Morozov, I.

2018-01-01

The paper presents the calculating results of the combustion products parameters in the tract of the low thrust rocket engine with thrust P ∼ 100 N. The article contains the following data: streamlines, distribution of total temperature parameter in the longitudinal section of the engine chamber, static temperature distribution in the cross section of the engine chamber, velocity distribution of the combustion products in the outlet section of the engine nozzle, static temperature near the inner wall of the engine. The presented parameters allow to estimate the efficiency of the mixture formation processes, flow of combustion products in the engine chamber and to estimate the thermal state of the structure.

Low-thrust chemical propulsion system propellant expulsion and thermal conditioning study. Executive summary

NASA Technical Reports Server (NTRS)

Merino, F.; Wakabayashi, I.; Pleasant, R. L.; Hill, M.

1982-01-01

Preferred techniques for providing abort pressurization and engine feed system net positive suction pressure (NPSP) for low thrust chemical propulsion systems (LTPS) were determined. A representative LTPS vehicle configuration is presented. Analysis tasks include: propellant heating analysis; pressurant requirements for abort propellant dump; and comparative analysis of pressurization techniques and thermal subcoolers.

Nuclear design of a vapor core reactor for space nuclear propulsion

NASA Astrophysics Data System (ADS)

Dugan, Edward T.; Watanabe, Yoichi; Kuras, Stephen A.; Maya, Isaac; Diaz, Nils J.

1993-01-01

Neutronic analysis methodology and results are presented for the nuclear design of a vapor core reactor for space nuclear propulsion. The Nuclear Vapor Thermal Reactor (NVTR) Rocket Engine uses modified NERVA geometry and systems which the solid fuel replaced by uranium tetrafluoride vapor. The NVTR is an intermediate term gas core thermal rocket engine with specific impulse in the range of 1000-1200 seconds; a thrust of 75,000 lbs for a hydrogen flow rate of 30 kg/s; average core exit temperatures of 3100 K to 3400 K; and reactor thermal powers of 1400 to 1800 MW. Initial calculations were performed on epithermal NVTRs using ZrC fuel elements. Studies are now directed at thermal NVTRs that use fuel elements made of C-C composite. The large ZrC-moderated reactors resulted in thrust-to-weight ratios of only 1 to 2; the compact C-C composite systems yield thrust-to-weight ratios of 3 to 5.

Low Thrust, Deep Throttling, US/CIS Integrated NTRE

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Kolganov, Vyacheslav; Rochow, Richard F.

1994-07-01

In 1993 our international team performed a follow-on ``Nuclear Thermal Rocket Engine (NTRE) Extended Life Feasibility Assessment'' study for the Nuclear Propulsion Office (NPO) at NASAs Lewis Research Center. The main purpose of this study was to complete the 1992 study matrix to assess NTRE designs at thrust levels of 22.5, 11.3, and 6.8 tonnes, using Commonwealth of Independent States (CIS) reactor technology. An additional Aerojet goal was to continue improving the NTRE concept we had generated. Deep throttling, mission performance optimized engine design parametrics, and reliability/cost enhancing engine system simplifications were studied, because they seem to be the last three basic design improvements sorely needed by post-NERVA NTRE. Deep throttling improves engine life by eliminating damaging thermal and mechanical shocks caused by after-cooling with pulsed coolant flow. Alternately, it improves mission performance with steady flow after-cooling by minimizing reactor over-cooling. Deep throttling also provides a practical transition from high pressures and powers of the high thrust power cycle to the low pressures and powers of our electric power generating mode. Two deep throttling designs are discussed; a workable system that was studied and a simplified system that is recommended for future study. Mission-optimized engine thrust/weight (T/W) and Isp predictions are included along with system flow schemes and concept sketches.

Thermal barrier coatings (TBC's) for high heat flux thrust chambers

NASA Astrophysics Data System (ADS)

Bradley, Christopher M.

The last 30 years materials engineers have been under continual pressure to develop materials with a greater temperature potential or to produce configurations that can be effectively cooled or otherwise protected at elevated temperature conditions. Turbines and thrust chambers produce some of the harshest service conditions for materials which lead to the challenges engineers face in order to increase the efficiencies of current technologies due to the energy crisis that the world is facing. The key tasks for the future of gas turbines are to increase overall efficiencies to meet energy demands of a growing world population and reduce the harmful emissions to protect the environment. Airfoils or blades tend to be the limiting factor when it comes to the performance of the turbine because of their complex design making them difficult to cool as well as limitations of their thermal properties. Key tasks for space transportation it to lower costs while increasing operational efficiency and reliability of our space launchers. The important factor to take into consideration is the rocket nozzle design. The design of the rocket nozzle or thrust chamber has to take into account many constraints including external loads, heat transfer, transients, and the fluid dynamics of expanded hot gases. Turbine engines can have increased efficiencies if the inlet temperature for combustion is higher, increased compressor capacity and lighter weight materials. In order to push for higher temperatures, engineers need to come up with a way to compensate for increased temperatures because material systems that are being used are either at or near their useful properties limit. Before thermal barrier coatings were applied to hot-section components, material alloy systems were able to withstand the service conditions necessary. But, with the increased demand for performance, higher temperatures and pressures have become too much for those alloy systems. Controlled chemistry of hot-section components has become critical, but at the same time the service conditions have put our best alloy systems to their limits. As a result, implementation of cooling holes and thermal barrier coatings are new advances in hot-section technologies now looked at for modifications to reach higher temperature applications. Current thermal barrier coatings used in today's turbine applications is known as 8%yttria-stabilized zirconia (YSZ) and there are no coatings for current thrust chambers. Current research is looking at the applicability of 8%yttria-stabilized hafnia (YSH) for turbine applications and the implementation of 8%YSZ onto thrust chambers. This study intends to determine if the use of thermal barrier coatings are applicable for high heat flux thrust chambers using industrial YSZ will be advantageous for improvements in efficiency, thrust and longer service life by allowing the thrust chambers to be used more than once.

Research and Technology

NASA Image and Video Library

1998-01-09

The Direct Gain Solar Thermal Engine was designed with no moving parts. The concept of Solar Thermal Propulsion Research uses focused solar energy from an inflatable concentrator (a giant magnifying glass) to heat a propellant (hydrogen) and allows thermal expansion through the nozzle for low thrust without chemical combustion. Energy limitations and propellant weight associated with traditional combustion engines are non-existant with this concept. The Direct Gain Solar Thermal Engine would be used for moving from a lower orbit to an upper synchronous orbit.

High heat transfer oxidizer heat exchanger design and analysis. [RL10-2B engine

NASA Technical Reports Server (NTRS)

Kmiec, Thomas D.; Kanic, Paul G.; Peckham, Richard J.

1987-01-01

The RL10-2B engine, a derivative of the RL10, is capable of multimode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2% of full thrust, and pumped idle (PI), which is 10% of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-2B engine during the low thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidizer heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. The design, concept verification testing and analysis for such a heat exchanger is discussed. The design presented uses a high efficiency compact core to vaporize the oxygen, and in the self-contained unit, attenuates any pressure and flow oscillations which result from unstable boiling in the core. This approach is referred to as the high heat transfer design. An alternative approach which prevents unstable boiling of the oxygen by limiting the heat transfer is referred to as the low heat transfer design and is reported in Pratt & Whitney report FR-19135-2.

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

1994-01-01

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

Gerrish, Harold P., Jr.

2003-01-01

This paper presents viewgraphs on Solar Thermal Propulsion (STP). Some of the topics include: 1) Ways to use Solar Energy for Propulsion; 2) Solar (fusion) Energy; 3) Operation in Orbit; 4) Propulsion Concepts; 5) Critical Equations; 6) Power Efficiency; 7) Major STP Projects; 8) Types of STP Engines; 9) Solar Thermal Propulsion Direct Gain Assembly; 10) Specific Impulse; 11) Thrust; 12) Temperature Distribution; 13) Pressure Loss; 14) Transient Startup; 15) Axial Heat Input; 16) Direct Gain Engine Design; 17) Direct Gain Engine Fabrication; 18) Solar Thermal Propulsion Direct Gain Components; 19) Solar Thermal Test Facility; and 20) Checkout Results.

Modified RS2101 rocket engine study program

NASA Technical Reports Server (NTRS)

1971-01-01

The purpose of the program is to perform design studies and analyses to determine the effects of incorporating a 60:1 expansion area ratio nozzle extension, extended firing time, and modified operating conditions and environments on the MM'71 rocket engine assembly. An injector-to-thrust chamber seal study was conducted to define potential solutions for leakage past this joint. The results and recommendations evolving from the engine thermal analyses, the injector-to-thrust chamber seal studies, and the nozzle extension joint stress analyses are presented.

A thermal control approach for a solar electric propulsion thrust subsystem

NASA Technical Reports Server (NTRS)

Maloy, J. E.; Oglebay, J. C.

1979-01-01

A thrust subsystem thermal control design is defined for a Solar Electric Propulsion System (SEPS) proposed for the comet Halley Flyby/comet Tempel 2 rendezvous mission. A 114 node analytic model, developed and coded on the systems improved numerical differencing analyzer program, was employed. A description of the resulting thrust subsystem thermal design is presented as well as a description of the analytic model and comparisons of the predicted temperature profiles for various SEPS thermal configurations that were generated using this model. It was concluded that: (1) a BIMOD engine system thermal design can be autonomous; (2) an independent thrust subsystem thermal design is feasible; (3) the interface module electronics temperatures can be controlled by a passive radiator and supplementary heaters; (4) maintaining heat pipes above the freezing point would require an additional 322 watts of supplementary heating power for the situation where no thrusters are operating; (5) insulation is required around the power processors, and between the interface module and the avionics module, as well as in those areas which may be subjected to solar heating; and (6) insulation behind the heat pipe radiators is not necessary.

Multiphysics Nuclear Thermal Rocket Thrust Chamber Analysis

NASA Technical Reports Server (NTRS)

Wang, Ten-See

2005-01-01

The objective of this effort is t o develop an efficient and accurate thermo-fluid computational methodology to predict environments for hypothetical thrust chamber design and analysis. The current task scope is to perform multidimensional, multiphysics analysis of thrust performance and heat transfer analysis for a hypothetical solid-core, nuclear thermal engine including thrust chamber and nozzle. The multiphysics aspects of the model include: real fluid dynamics, chemical reactivity, turbulent flow, and conjugate heat transfer. The model will be designed to identify thermal, fluid, and hydrogen environments in all flow paths and materials. This model would then be used to perform non- nuclear reproduction of the flow element failures demonstrated in the Rover/NERVA testing, investigate performance of specific configurations and assess potential issues and enhancements. A two-pronged approach will be employed in this effort: a detailed analysis of a multi-channel, flow-element, and global modeling of the entire thrust chamber assembly with a porosity modeling technique. It is expected that the detailed analysis of a single flow element would provide detailed fluid, thermal, and hydrogen environments for stress analysis, while the global thrust chamber assembly analysis would promote understanding of the effects of hydrogen dissociation and heat transfer on thrust performance. These modeling activities will be validated as much as possible by testing performed by other related efforts.

Nuclear Engine System Simulation (NESS) version 2.0

NASA Technical Reports Server (NTRS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.

1993-01-01

The topics are presented in viewgraph form and include the following; nuclear thermal propulsion (NTP) engine system analysis program development; nuclear thermal propulsion engine analysis capability requirements; team resources used to support NESS development; expanded liquid engine simulations (ELES) computer model; ELES verification examples; NESS program development evolution; past NTP ELES analysis code modifications and verifications; general NTP engine system features modeled by NESS; representative NTP expander, gas generator, and bleed engine system cycles modeled by NESS; NESS program overview; NESS program flow logic; enabler (NERVA type) nuclear thermal rocket engine; prismatic fuel elements and supports; reactor fuel and support element parameters; reactor parameters as a function of thrust level; internal shield sizing; and reactor thermal model.

Revised Point of Departure Design Options for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Fittje, James E.; Borowski, Stanley K.; Schnitzler, Bruce

2015-01-01

In an effort to further refine potential point of departure nuclear thermal rocket engine designs, four proposed engine designs representing two thrust classes and utilizing two different fuel matrix types are designed and analyzed from both a neutronics and thermodynamic cycle perspective. Two of these nuclear rocket engine designs employ a tungsten and uranium dioxide cermet (ceramic-metal) fuel with a prismatic geometry based on the ANL-200 and the GE-710, while the other two designs utilize uranium-zirconium-carbide in a graphite composite fuel and a prismatic fuel element geometry developed during the Rover/NERVA Programs. Two engines are analyzed for each fuel type, a small criticality limited design and a 111 kN (25 klbf) thrust class engine design, which has been the focus of numerous manned mission studies, including NASA's Design Reference Architecture 5.0. slightly higher T/W ratios, but they required substantially more 235U.

Oxidizer heat exchanger component test

NASA Technical Reports Server (NTRS)

Kanic, P. G.

1988-01-01

The RL10-IIB engine, is capable of multimode thrust operation. The engine operates at two low-thrust levels: tank head idle (THI), approximately 1 to 2 percent of full thrust; and pumped idle, 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient thermal conditioning; PI operation provides vehicle tank prepressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-IIB engine during the low-thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidized heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. This report summarizes the test activity and post-test data analysis for two possible heat exchangers, each of which employs a completely different design philosophy. One design makes use of a low-heat transfer (PHT) approach in combination with a volume to attenuate pressure and flow oscillations. The test data showed that the LHT unit satisfied the oxygen exit quality of 0.95 or greater in both the THI and PI modes while maintaining stability. The HHT unit fulfilled all PI requirements; data for THI satisfactory operation is implied from experimental data that straddle the exact THI operating point.

Research Technology

NASA Image and Video Library

1999-11-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. This photograph shows a fully assembled solar thermal engine placed inside the vacuum chamber at the test facility prior to testing. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move theNation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Solar Thermal Propulsion Test Facility

NASA Technical Reports Server (NTRS)

1999-01-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. This photograph shows a fully assembled solar thermal engine placed inside the vacuum chamber at the test facility prior to testing. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move theNation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

ESCORT: A Pratt & Whitney nuclear thermal propulsion and power system for manned mars missions

NASA Astrophysics Data System (ADS)

Feller, Gerald J.; Joyner, Russell

1999-01-01

The purpose of this paper is to describe the conceptual design of an upgrade to the Pratt & Whitney ESCORT nuclear thermal rocket engine. The ESCORT is a bimodal engine capable of supporting a wide range of vehicle propulsive and electrical power requirements. The ESCORT engine is powered by a fast-spectrum beryllium-reflected CERMET-fueled nuclear reactor. In propulsive mode, the reactor is used to heat hot hydrogen to approximately 2700 K which is expanded through a converging/diverging nozzle to generate thrust. Heat pickup in the nozzle and the radial beryllium reflectors is used to drive the turbomachinery in the ESCORT expander cycle. In electrical mode, the reactor is used to heat a mixture of helium and xenon to drive a closed-loop Brayton cycle in order to generate electrical energy. This closed loop system has the additional function of a decay heat removal system after the propulsive mode operation is discontinued. The original ESCORT design was capable of delivering 4448.2 N (1000 lbf) of thrust at a vacuum impulse level of approximately 900 s. Design Reference Mission requirements (DRM) from NASA Johnson Space Center and NASA Lewis Research Center studies in 1997 and 1998 have detailed upgraded requirements for potential manned Mars missions. The current NASA DRM requires a nuclear thermal propulsion system capable of delivering total mission requirements of 200170 N (45000 lbf) thrust and 50 kWe of spacecraft electrical power. This is met assuming three engines capable of each delivering 66723 N (15000 lbf) of vacuum thrust and 25 kWe of electrical power. The individual engine requirements were developed assuming three out of three engine reliability for propulsion and two out of three engine reliability for spacecraft electrical power. The approximate target vacuum impulse is 925 s. The Pratt & Whitney ESCORT concept was upgraded to meet these requirements. The hexagonal prismatic fuel elements were modified to address the uprated power requirements while maintaining the peak fuel temperature below the 2880 K limit for W-UO2 CERMET fuels. A system integrated performance methodology was developed to assess the sensitivity to weight, thrust and impulse to the DRM requirements. Propellant tanks, shielding, and Brayton cycle power conversion unit requirements were included in this evaluation.

Solar Thermal Propulsion Test Facility at MSFC

NASA Technical Reports Server (NTRS)

1999-01-01

This photograph shows an overall view of the Solar Thermal Propulsion Test Facility at the Marshall Space Flight Center (MSFC). The 20-by 24-ft heliostat mirror, shown at the left, has dual-axis control that keeps a reflection of the sunlight on an 18-ft diameter concentrator mirror (right). The concentrator mirror then focuses the sunlight to a 4-in focal point inside the vacuum chamber, shown at the front of concentrator mirror. Researchers at MSFC have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than chemical a combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propell nt. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Non-Toxic Dual Thrust Reaction Control Engine Development for On-Orbit APS Applications

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.

2003-01-01

A non-toxic dual thrust proof-of-concept demonstration engine was successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the Next Generation Launch Technology (NGLT) program. The demonstration engine utilized the existing Kistler K-1 870 lbf LOX/Ethanol orbital maneuvering engine ( O m ) coupled with some special test equipment (STE) that enabled engine operation at 870 lbf in the primary mode and 25 lbf in the vernier mode. Ambient testing in primary mode varied mixture ratio (MR) from 1.28 to 1.71 and chamber pressure (P(c) from 110 to 181 psia, and evaluated electrical pulse widths (EPW) of 0.080, 0.100 and 0.250 seconds. Altitude testing in vernier mode explored igniter and thruster pulsing characteristics, long duration steady state operation (greater than 420 sec) and the impact of varying the percent fuel film cooling on vernier performance and chamber thermal response at low PC (4 psia). Data produced from the testing provided calibration of the performance and thermal models used in the design of the next version of the dual thrust Reaction Control Engine (RCE).

Breadboard RL10-2B low-thrust operating mode (second iteration) test report

NASA Technical Reports Server (NTRS)

Kanic, Paul G.; Kaldor, Raymond B.; Watkins, Pia M.

1988-01-01

Cryogenic rocket engines requiring a cooling process to thermally condition the engine to operating temperature can be made more efficient if cooling propellants can be burned. Tank head idle and pumped idle modes can be used to burn propellants employed for cooling, thereby providing useful thrust. Such idle modes required the use of a heat exchanger to vaporize oxygen prior to injection into the combustion chamber. During December 1988, Pratt and Whitney conducted a series of engine hot firing demonstrating the operation of two new, previously untested oxidizer heat exchanger designs. The program was a second iteration of previous low thrust testing conducted in 1984, during which a first-generation heat exchanger design was used. Although operation was demonstrated at tank head idle and pumped idle, the engine experienced instability when propellants could not be supplied to the heat exchanger at design conditions.

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Alternate low-thrust capability task report

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1980-01-01

The feasibility and design impact of a requirement for the advanced expander cycle engine to be adaptable to extended low thrust operation of approximately 1K to 2K lb is assessed. It is determined that the orbit transfer vehicle point design engine can be reduced in thrust with minor injector modifications from 15K to 1K without significantly affecting combustion performance efficiency or injector face/chamber wall thermal compatibility. Likewise, high frequency transverse mode combustion instability is not expected to be detrimentally affected. Primarily, the operational limitations consist of feed system chugging instabilities and potential coupling of the injector response with the chamber longitudinal mode resonances under certain operating conditions. The recommended injector modification for low thrust operation is a change in the oxidizer injector element orifice size. Analyses also indicate that chamber coolant flow stability may be a concern below 2K 1bF operation and oxidizer pump stability could be a problem below a 2K thrust level although a recirculation flow could alleviate the problem.

Review of Nuclear Thermal Propulsion Ground Test Options

NASA Technical Reports Server (NTRS)

Coote, David J.; Power, Kevin P.; Gerrish, Harold P.; Doughty, Glen

2015-01-01

High efficiency rocket propulsion systems are essential for humanity to venture beyond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rockets with relatively high thrust and twice the efficiency of highest performing chemical propellant engines. NTP utilizes the coolant of a nuclear reactor to produce propulsive thrust. An NTP engine produces thrust by flowing hydrogen through a nuclear reactor to cool the reactor, heating the hydrogen and expelling it through a rocket nozzle. The hot gaseous hydrogen is nominally expected to be free of radioactive byproducts from the nuclear reactor; however, it has the potential to be contaminated due to off-nominal engine reactor performance. NTP ground testing is more difficult than chemical engine testing since current environmental regulations do not allow/permit open air testing of NTP as was done in the 1960's and 1970's for the Rover/NERVA program. A new and innovative approach to rocket engine ground test is required to mitigate the unique health and safety risks associated with the potential entrainment of radioactive waste from the NTP engine reactor core into the engine exhaust. Several studies have been conducted since the ROVER/NERVA program in the 1970's investigating NTP engine ground test options to understand the technical feasibility, identify technical challenges and associated risks and provide rough order of magnitude cost estimates for facility development and test operations. The options can be divided into two distinct schemes; (1) real-time filtering of the engine exhaust and its release to the environment or (2) capture and storage of engine exhaust for subsequent processing.

LANTR Engine Optimization for Lunar Missions

NASA Astrophysics Data System (ADS)

Bulman, M. J.; Poth, Greg; Borowski, Stan

2006-01-01

Propulsion requirements for sustainable Lunar missions are very demanding. The high Delta V for short transit times and/or reusable vehicles are best served with the High Isp of Nuclear Propulsion. High thrust is needed to reduce gravity losses during earth departure. The LOX-Augmented Nuclear Thermal Rocket (LANTR) is a concept whereby thrust from a nuclear thermal rocket can be doubled, or even quadrupled, by the injection and combustion of gaseous oxygen downstream of the throat. This has many advantages for the mission including a reduction in the size of the reactor(s) and propellant tank volume for a given payload delivered to Low Lunar Orbit. In this paper, we conduct mission studies to define the optimum basic (Unaugmented) engine thrust, Lox augmentation level and Lox loading for minimum initial mass in low earth orbit. 35% mass savings are seen for NTR powered LTVs with over twice the propellant Volume. The LANTR powered LTV has a similar mass savings with minimal volume penalties.

Cyclic hot firing results of tungsten-wire-reinforced, copper-lined thrust chambers

NASA Technical Reports Server (NTRS)

Kazaroff, John M.; Jankovsky, Robert S.

1990-01-01

An advanced thrust liner material for potential long life reusable rocket engines is described. This liner material was produced with the intent of improving the reusable life of high pressure thrust chambers by strengthening the chamber in the hoop direction, thus avoiding the longitudinal cracking due to low cycle fatigue that is observed in conventional homogeneous copper chambers, but yet not reducing the high thermal conductivity that is essential when operating with high heat fluxes. The liner material produced was a tungsten wire reinforced copper composite. Incorporating this composite into two hydrogen-oxygen test rocket chambers was done so that its performance as a reusable liner material could be evaluated. Testing results showed that both chambers failed prematurely, but the crack sites were perpendicular to the normal direction of cracking indicating a degree of success in containing the tremendous thermal strain associated with high temperature rocket engines. The failures, in all cases, were associated with drilled instrumentation ports and no other damages or deformations were found elsewhere in the composite liners.

The research Of Multilayer Thermal Insulation With Mechanical Properties Based On Model Analysis Test

NASA Astrophysics Data System (ADS)

Lianhua, Yin

The heat shield of aircraft is made of the major thrusts structure with multilayer thermal insulation part. For protecting against thermo-radiation from larger thrusting force engine,the heat shield is installed around this engine nearby.The multilayer thermal insulation part with multilayer radiation/reflection structure is made of reflection layer and interval layer.At vacuum condition,these materials is higher heat insulation capability than other material,is applied for lots of pats on aircraft extensively.But because of these material is made of metal and nonmetal,it is impossible to receive it's mechanical properties of materials from mechanical tests.These paper describes a new measure of mechanical properties of materials in the heat shield based on model analysis test.At the requirement for the first order lateral frequency,these measure provide for the FEM analysis foundation on the optimization structure of the heat shield.

Perspective on thermal barrier coatings for industrial gas turbine applications

NASA Technical Reports Server (NTRS)

Mutasim, Zaher; Brentnall, William

1995-01-01

Thermal barrier coatings (TBC's) have been used in high thrust aircraft engines for many years, and have proved to be very effective in providing thermal protection and increasing engine efficiencies. TBC life requirements for aircraft engines are typically less than those required for industrial gas turbines. This paper describes current and future applications of TBC's in industrial gas turbine engines. Early testing and applications of TBC's is reviewed. Areas of concern from the engine designer's and materials engineer's perspective are identified and evaluated. This paper focuses on the key factors that are expected to influence utilization of TBC's in advanced industrial gas turbine engines. It is anticipated that reliable, durable and highly effective coating systems will be produced that will ultimately improve engine efficiency and performance.

Performance Capability of Single-Cavity Vortex Gaseous Nuclear Rockets

NASA Technical Reports Server (NTRS)

Ragsdale, Robert G.

1963-01-01

An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher hydrogen flow rates offer the possibility of sufficiently high thrust-to-weight ratios, if they are not limited by short thermal-radiation path lengths.

Solar Thermal Propulsion Test

NASA Technical Reports Server (NTRS)

1999-01-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. This image, taken during the test, depicts the light being concentrated into the focal point inside the vacuum chamber. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Research Technology

NASA Image and Video Library

1999-11-01

This photograph shows an overall view of the Solar Thermal Propulsion Test Facility at the Marshall Space Flight Center (MSFC). The 20-by 24-ft heliostat mirror, shown at the left, has dual-axis control that keeps a reflection of the sunlight on an 18-ft diameter concentrator mirror (right). The concentrator mirror then focuses the sunlight to a 4-in focal point inside the vacuum chamber, shown at the front of concentrator mirror. Researchers at MSFC have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than chemical a combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propell nt. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Characterization of a Pressure-Fed LOX/LCH4 Reaction Control System Under Simulated Altitude and Thermal Vacuum Conditions

NASA Technical Reports Server (NTRS)

Atwell, Matthew J.; Melcher, John C.; Hurlbert, Eric A.; Morehead, Robert L.

2017-01-01

A liquid oxygen, liquid methane (LOX/LCH4) reaction control system (RCS) was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under simulated altitude and thermal vacuum conditions. The RCS is a subsystem of the Integrated Cryogenic Propulsion Test Article (ICPTA) and was initially developed under Project Morpheus. Composed of two 28 lbf-thrust and two 7 lbf-thrust engines, the RCS is fed in parallel with the ICPTA main engine from four propellant tanks. 40 tests consisting of 1,010 individual thruster pulses were performed across 6 different test days. Major test objectives were focused on system dynamics, and included characterization of fluid transients, manifold priming, manifold thermal conditioning, thermodynamic vent system (TVS) performance, and main engine/RCS interaction. Peak surge pressures from valve opening and closing events were examined. It was determined that these events were impacted significantly by vapor cavity formation and collapse. In most cases the valve opening transient was more severe than the valve closing. Under thermal vacuum conditions it was shown that TVS operation is unnecessary to maintain liquid conditions at the thruster inlets. However, under higher heat leak environments the RCS can still be operated in a self-conditioning mode without overboard TVS venting, contingent upon the engines managing a range of potentially severe thermal transients. Lastly, during testing under cold thermal conditions the engines experienced significant ignition problems. Only after warming the thruster bodies with a gaseous nitrogen purge to an intermediate temperature was successful ignition demonstrated.

Focal Point Inside the Vacuum Chamber for Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

1999-01-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. The 20- by 24-ft heliostat mirror (not shown in this photograph) has dual-axis control that keeps a reflection of the sunlight on an 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. This photograph is a close-up view of a 4-in focal point inside the vacuum chamber at the MSFC Solar Thermal Propulsion Test facility. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Investigation of a Tricarbide Grooved Ring Fuel Element for a Nuclear Thermal Rocket

NASA Technical Reports Server (NTRS)

Taylor, Brian D.; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2017-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the temperature limitations of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment.

Solar Thermal Propulsion Test Facility

NASA Technical Reports Server (NTRS)

1999-01-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. This photograph, taken at MSFC's Solar Thermal Propulsion Test Facility, shows a concentrator mirror, a combination of 144 mirrors forming this 18-ft diameter concentrator, and a vacuum chamber that houses the focal point. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-foot diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Parametric scramjet analysis

NASA Astrophysics Data System (ADS)

Choi, Jongseong

The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a scramjet propulsion system to show how they relate to scramjet performance and the seven scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.

SSME structural computer program development. Volume 3: BOPACE demonstration analysis of the SSME thrust chamber liner

NASA Technical Reports Server (NTRS)

Spring, A. H.

1973-01-01

The application of a structural computer program for analysis of a thrust chamber liner is discussed. Two objectives were accomplished as follows: (1) exercise of the full capabilities of the computer program and (2) definition of thermal and mechanical boundary conditions to reflect the emergency power level operating conditions for the SSME 47OK engine at a station just upstream of the thrust chamber throat. Creep information on the thrust chamber is presented as a reference curve of creep strain versus time for various temperatures. Contour plots of the effective plastic strain, effective stress, and effective creep strain are developed.

A Crewed Mission to Apophis Using a Hybrid Bimodal Nuclear Thermal Electric Propulsion (BNTEP) System

NASA Technical Reports Server (NTRS)

Mccurdy, David R.; Borowski, Stanley K.; Burke, Laura M.; Packard, Thomas W.

2014-01-01

A BNTEP system is a dual propellant, hybrid propulsion concept that utilizes Bimodal Nuclear Thermal Rocket (BNTR) propulsion during high thrust operations, providing 10's of kilo-Newtons of thrust per engine at a high specific impulse (Isp) of 900 s, and an Electric Propulsion (EP) system during low thrust operations at even higher Isp of around 3000 s. Electrical power for the EP system is provided by the BNTR engines in combination with a Brayton Power Conversion (BPC) closed loop system, which can provide electrical power on the order of 100's of kWe. High thrust BNTR operation uses liquid hydrogen (LH2) as reactor coolant propellant expelled out a nozzle, while low thrust EP uses high pressure xenon expelled by an electric grid. By utilizing an optimized combination of low and high thrust propulsion, significant mass savings over a conventional NTR vehicle can be realized. Low thrust mission events, such as midcourse corrections (MCC), tank settling burns, some reaction control system (RCS) burns, and even a small portion at the end of the departure burn can be performed with EP. Crewed and robotic deep space missions to a near Earth asteroid (NEA) are best suited for this hybrid propulsion approach. For these mission scenarios, the Earth return V is typically small enough that EP alone is sufficient. A crewed mission to the NEA Apophis in the year 2028 with an expendable BNTEP transfer vehicle is presented. Assembly operations, launch element masses, and other key characteristics of the vehicle are described. A comparison with a conventional NTR vehicle performing the same mission is also provided. Finally, reusability of the BNTEP transfer vehicle is explored.

Space transportation booster engine thrust chamber technology, large scale injector

NASA Technical Reports Server (NTRS)

Schneider, J. A.

1993-01-01

The objective of the Large Scale Injector (LSI) program was to deliver a 21 inch diameter, 600,000 lbf thrust class injector to NASA/MSFC for hot fire testing. The hot fire test program would demonstrate the feasibility and integrity of the full scale injector, including combustion stability, chamber wall compatibility (thermal management), and injector performance. The 21 inch diameter injector was delivered in September of 1991.

Parametric analysis of a down-scaled turbo jet engine suitable for drone and UAV propulsion

NASA Astrophysics Data System (ADS)

Wessley, G. Jims John; Chauhan, Swati

2018-04-01

This paper presents a detailed study on the need for downscaling gas turbine engines for UAV and drone propulsion. Also, the procedure for downscaling and the parametric analysis of a downscaled engine using Gas Turbine Simulation Program software GSP 11 is presented. The need for identifying a micro gas turbine engine in the thrust range of 0.13 to 4.45 kN to power UAVs and drones weighing in the range of 4.5 to 25 kg is considered and in order to meet the requirement a parametric analysis on the scaled down Allison J33-A-35 Turbojet engine is performed. It is evident from the analysis that the thrust developed by the scaled engine and the Thrust Specific Fuel Consumption TSFC depends on pressure ratio, mass flow rate of air and Mach number. A scaling factor of 0.195 corresponding to air mass flow rate of 7.69 kg/s produces a thrust in the range of 4.57 to 5.6 kN while operating at a Mach number of 0.3 within the altitude of 5000 to 9000 m. The thermal and overall efficiency of the scaled engine is found to be 67% and 75% respectively for a pressure ratio of 2. The outcomes of this analysis form a strong base for further analysis, design and fabrication of micro gas turbine engines to propel future UAVs and drones.

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

NASA Technical Reports Server (NTRS)

Burke, Laura M.; Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2013-01-01

A crewed mission to Mars poses a significant challenge in dealing with the physiological issues that arise with the crew being exposed to a near zero-gravity environment as well as significant solar and galactic radiation for such a long duration. While long surface stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological effects on the crew. However, for a 1-year round trip mission, the outbound and inbound hyperbolic velocity at Earth and Mars can be very large resulting in a significant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power levels (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower specific mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for efficient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo-Newtons of thrust at reasonably high specific impulse (Isp) of 900 seconds for impulsive transplanetary injection and orbital insertion maneuvers. When in power generation/EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

NASA Technical Reports Server (NTRS)

Burke, Laura A.; Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2013-01-01

A crewed mission to Mars poses a signi cant challenge in dealing with the physiolog- ical issues that arise with the crew being exposed to a near zero-gravity environment as well as signi cant solar and galactic radiation for such a long duration. While long sur- face stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological e ects on the crew. However, for a 1-year round trip mission, the outbound and inbound hy- perbolic velocity at Earth and Mars can be very large resulting in a signi cant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power lev- els (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower speci c mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for ecient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo- Newtons of thrust at reasonably high speci c impulse (Isp) of 900 seconds for impulsive trans-planetary injection and orbital insertion maneuvers. When in power generation / EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Research Technology

NASA Image and Video Library

1999-08-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. The 20- by 24-ft heliostat mirror (not shown in this photograph) has dual-axis control that keeps a reflection of the sunlight on an 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. This photograph is a close-up view of a 4-in focal point inside the vacuum chamber at the MSFC Solar Thermal Propulsion Test facility. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Research Technology

NASA Image and Video Library

1999-08-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. This photograph, taken at MSFC's Solar Thermal Propulsion Test Facility, shows a concentrator mirror, a combination of 144 mirrors forming this 18-ft diameter concentrator, and a vacuum chamber that houses the focal point. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-foot diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Performance Evaluation of the T6 Ion Engine

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

2010-01-01

The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

Research Technology

NASA Image and Video Library

1999-03-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. Solar thermal propulsion works by acquiring and redirecting solar energy to heat a propellant. The 20- by 24-ft heliostat mirror (not shown in this photograph) has a dual-axis control that keeps a reflection of the sunlight on the 18-ft diameter concentrator mirror, which then focuses the sunlight to a 4-in focal point inside the vacuum chamber. The focal point has 10 kilowatts of intense solar power. This image, taken during the test, depicts the light being concentrated into the focal point inside the vacuum chamber. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth orbit, rapid travel throughout the solar system, and exploration of interstellar space.

Clustered engine study

NASA Technical Reports Server (NTRS)

Shepard, Kyle; Sager, Paul; Kusunoki, Sid; Porter, John; Campion, AL; Mouritzan, Gunnar; Glunt, George; Vegter, George; Koontz, Rob

1993-01-01

Several topics are presented in viewgraph form which together encompass the preliminary assessment of nuclear thermal rocket engine clustering. The study objectives, schedule, flow, and groundrules are covered. This is followed by the NASA groundrules mission and our interpretation of the associated operational scenario. The NASA reference vehicle is illustrated, then the four propulsion system options are examined. Each propulsion system's preliminary design, fluid systems, operating characteristics, thrust structure, dimensions, and mass properties are detailed as well as the associated key propulsion system/vehicle interfaces. A brief series of systems analysis is also covered including: thrust vector control requirements, engine out possibilities, propulsion system failure modes, surviving system requirements, and technology requirements. An assessment of vehicle/propulsion system impacts due to the lessons learned are presented.

Experience with high performance V/STOL fighter projects at MBB

NASA Technical Reports Server (NTRS)

Aulehla, F.; Kissel, G. K.

1981-01-01

Flight control systems and aerodynamic aspects of experimental V/STOL aircraft are discussed. The VJ 101 C featured tilting engines for increased thrust, reheat for takeoff, simple translation, triangular decentralization of the engines for thrust modulation, and moderate ground effects. Two experimental aircraft were built, with and without reheat, capable of Mach 2 and Mach 1.04, respectively. The mechanical flight control system and tests are outlined, both for hover rig and flight configurations. Ground suction, acoustic and thermal loading, sodium silicate coatings to avoid ground corrosion, and recirculation are considered. Results of the follow-on project to the VJ 101 C, the AVS, which was developed by NASA, are reviewed, and it is noted that trends toward thrust-to-weight ratios exceeding one, in concert with low wing loading, favor the development of V/STOL aircraft.

The ENABLER - Based on proven NERVA technology

NASA Astrophysics Data System (ADS)

Livingston, Julie M.; Pierce, Bill L.

The ENABLER reactor for use in a nuclear thermal propulsion engine uses the technology developed in the NERVA/Rover program, updated to incorporate advances in the technology. Using composite fuel, higher power densities per fuel element, improved radiation resistant control components and the advancements in use of carbon-carbon materials; the ENABLER can provide a specific impulse of 925 seconds, an engine thrust to weight (excluding reactor shield) approaching five, an improved initial mass in low Earth orbit and a consequent reduction in launch costs and logistics problems. This paper describes the 75,000 lbs thrust ENABLER design which is a low cost, low risk approach to meeting tommorrow's space propulsion needs.

Affordable Development and Demonstration of a Small NTR engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Sefcik, R. J.; Fittje, J. E.; McCurdy, D. R.; Qualls, A. L.; Schnitzler, B. G; Werner, J.; Weitzberg, A.; Joyner, C. R.

2015-01-01

In FY'11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program A strategy was outlined by GRC and NASA HQ that included 2 key elements -"Foundational Technology Development" followed by specific "Technology Demonstration" projects. The "Technology Demonstration "element proposed ground technology demonstration (GTD) testing in the early 2020's, followed by a flight technology demonstration (FTD) mission by approx. 2025. In order to reduce development costs, the demonstration projects would focus on developing a small, low thrust (approx. 7.5 -16.5 klb(f)) engine that utilizes a "common" fuel element design scalable to the higher thrust (approx. 25 klb(f)) engines used in NASA's Mars DRA 5.0 study(NASA-SP-2009-566). Besides reducing development costs and allowing utilization of existing, flight proven engine hard-ware (e.g., hydrogen pumps and nozzles), small, lower thrust ground and flight demonstration engines can validate the technology and offer improved capability -increased payloads and decreased transit times -valued for robotic science missions identified in NASA's Decadal Study.

Analysis of a Nuclear Enhanced Airbreathing Rocket for Earth to Orbit Applications

NASA Technical Reports Server (NTRS)

Adams, Robert B.; Landrum, D. Brian; Brown, Norman (Technical Monitor)

2001-01-01

The proposed engine concept is the Nuclear Enhanced Airbreathing Rocket (NEAR). The NEAR concept uses a fission reactor to thermally heat a propellant in a rocket plenum. The rocket is shrouded, thus the exhaust mixes with ingested air to provide additional thermal energy through combustion. The combusted flow is then expanded through a nozzle to provide thrust.

Thermal-environmental testing of a 30-cm engineering model thruster

NASA Technical Reports Server (NTRS)

Mirtich, M. J.

1976-01-01

An experimental test program was carried out to document all 30-cm electron bombardment Hg ion bombardment thruster functions and characteristics over the thermal environment of several proposed missions. An engineering model thruster was placed in a thermal test facility equipped with -196 C walls and solar simulation. The thruster was cold soaked and exposed to simulated eclipses lasting in duration from 17 to 72 minutes. The thruster was operated at quarter, to full beam power in various thermal configurations which simulated multiple thruster operation, and was also exposed to 1 and 2 suns solar simulation. Thruster control characteristics and constraints; performance, including thrust magnitude and direction; and structural integrity were evaluated over the range of thermal environments tested.

Thermal-environment testing of a 30-cm engineering model thruster

NASA Technical Reports Server (NTRS)

Mirtich, M. J.

1976-01-01

An experimental test program was carried out to document all 30-cm electron bombardment Hg ion bombardment thruster functions and characteristics over the thermal environment of several proposed missions. An engineering model thruster was placed in a thermal test facility equipped with -196 C walls and solar simulation. The thruster was cold soaked and exposed to simulated eclipses lasting in duration from 17 to 72 minutes. The thruster was operated at quarter, to full beam power in various thermal configurations which simulated multiple thruster operation, and was also exposed to 1 and 2 suns solar simulation. Thruster control characteristics and constraints; performance, including thrust magnitude and direction; and structural integrity were evaluated over the range of thermal environments tested.

Earth storable bimodal engine, phase 1

NASA Technical Reports Server (NTRS)

1973-01-01

An in-depth study of an Earth Storable Bimodal (ESB) Engine using earth storable propellants N2O/N2H4 and operating in either a monopropellant or bipropellant mode was conducted. Detailed studies were completed for both a hot-gas, regeneratively cooled thrust chamber and a ducted hot-gas, film cooled thrust chamber. Hydrazine decomposition products were used for cooling in either configuration. The various arrangements and configurations of hydrazine reactors, secondary injectors, chambers and gimbal methods were considered. The two basic materials selected for the major components were columbium alloys and L-605. The secondary injector types considered were previously demonstrated by JPL and consisted of a liquid-on-gas triplet, a liquid-on-gas doublet, and a liquid-on-gas coaxial injector. Various design tradeoffs were made with different reactor types located at: the secondary injector station, the thrust chamber throat, and the nozzle/extension interface. Associated thermal, structural, and mass analyses were completed.

Adaptation Method for Overall and Local Performances of Gas Turbine Engine Model

NASA Astrophysics Data System (ADS)

Kim, Sangjo; Kim, Kuisoon; Son, Changmin

2018-04-01

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

The ENABLER—based on proven NERVA technology

NASA Astrophysics Data System (ADS)

Livingston, Julie M.; Pierce, Bill L.

1991-01-01

The ENABLER reactor for use in a nuclear thermal propulsion engine uses the technology developed in the NERVA/Rover program, updated to incorporate advances in the technology. Using composite fuel, higher power densities per fuel element, improved radiation resistant control components and the advancements in use of carbon-carbon materials; the ENABLER can provide a specific impulse of 925 seconds, an engine thrust to weight (excluding reactor shield) approaching five, an improved initial Mass In Low Earth Orbit (IMLEO) and a consequent reduction in launch costs and logistics problems. This paper describes the 75,000 lbs thrust ENABLER design which is a low cost, low risk approach to meeting tomorrow's space propulsion needs.

A first scramjet study

NASA Technical Reports Server (NTRS)

Emanuel, George

1989-01-01

A variety of related scramjet engine topics are examined. The flow is assumed to be 1-D, the gas is thermally and calorically perfect, and focus is on low hypersonic Mach numbers. The thrust and lift of an exposed half nozzle, which is used on the aerospace plane, is evaluated as well as a fully confined nozzle. A rough estimate of the drag of an aerospace plane is provided. Thermal effects and shock waves are next discussed. A parametric scramjet model is then presented based on the influence coefficient method, which evaluates the dominant scramjet processes. The independent parameters are the ratio of specific heats, a nondimensional heat addition parameter, and four Mach numbers. The total thrust generated by the combustor and nozzle is shown to be independent of the heat release distribution and the combustor exit Mach number, providing thermal choking is avoided. An operating condition for the combustor is found that maximizes the thrust. An alternative condition is explored when this optimum is no longer realistic. This condition provides a favorable pressure gradient and a reasonable area ratio for the combustor. Parametric results based on the model is provided.

Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster (NEXT)

NASA Technical Reports Server (NTRS)

Diamant, Kevin D.; Pollard, James E.; Crofton, Mark W.; Patterson, Michael J.; Soulas, George C.

2010-01-01

Direct thrust measurements have been made on the NASA Evolutionary Xenon Thruster (NEXT) ion engine using a standard pendulum style thrust stand constructed specifically for this application. Values have been obtained for the full 40-level throttle table, as well as for a few off-nominal operating conditions. Measurements differ from the nominal NASA throttle table 10 (TT10) values by 3.1 percent at most, while at 30 throttle levels (TLs) the difference is less than 2.0 percent. When measurements are compared to TT10 values that have been corrected using ion beam current density and charge state data obtained at The Aerospace Corporation, they differ by 1.2 percent at most, and by 1.0 percent or less at 37 TLs. Thrust correction factors calculated from direct thrust measurements and from The Aerospace Corporation s plume data agree to within measurement error for all but one TL. Thrust due to cold flow and "discharge only" operation has been measured, and analytical expressions are presented which accurately predict thrust based on thermal thrust generation mechanisms.

Extended performance solar electric propulsion thrust system study. Volume 2: Baseline thrust system

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Hawthorne, E. I.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30- cm engineering model thruster as the technology base. Emphasis was placed on relatively high-power missions (60 to 100 kW) such as a Halley's comet rendezvous. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power-processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. The baseline thrust system design features modular construction, conventional power processing, and a concentractor solar array concept and is designed to interface with the space shuttle.

Revised Point of Departure Design Options for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Fittje, James E.; Schnitzler, Bruce G.; Borowoski, Stanley

2015-01-01

Four Revised Point of Departure NTR Engines were Designed and Analyzed using MCNP and NESS. All Four Engines Have Thermodynamically Closed Cycles at Nominal Chamber Pressures. 111 kilonewton (25 kip-force) Cermet Design Required Dedicated Heater Elements to Close the Cycle. Cermet Based Designs had Slightly Higher TW Ratios, but Required Substantially More U-235. NERVA Derived Criticality Limited Engine Could Operate at Lower Power and Thrust Levels Compared to the Criticality Limited Cermet Design.

Extended performance solar electric propulsion thrust system study. Volume 3: Tradeoff studies of alternate thrust system configurations

NASA Technical Reports Server (NTRS)

Hawthorne, E. I.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. Emphasis was placed on relatively high power missions. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power-processing components were performed. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. A program development plan was formulated that outlines the work structure considered necessary for developing, qualifying, and fabricating the flight hardware for the baseline thrust system within the time frame of a project to rendezvous with Halley's comet. An assessment was made of the costs and risks associated with a baseline thrust system as provided to the mission project under this plan. Critical procurements and interfaces were identified and defined.

Perspective on thermal barrier coatings for industrial gas turbine applications

NASA Technical Reports Server (NTRS)

Mutasim, Z. Z.; Hsu, L. L.; Brentnall, W. D.

1995-01-01

Thermal Barrier Coatings (TBC's) have been used in high thrust aircraft engines for many years, and have proved to be very effective in allowing higher turbine inlet temperatures. TBC life requirements for aircraft engines are typically less than those required in industrial gas turbines. The use of TBC's for industrial gas turbines can increase if durability and longer service life can be successfully demonstrated. This paper will describe current and future applications of TBC's in industrial gas turbine engines. Early testing and applications of TBC's will also be reviewed. This paper focuses on the key factors that are expected to influence utilization of TBC's in advanced industrial gas turbine engines. It is anticipated that reliable, durable and high effective coating systems will be produced that will ultimately improve engine efficiency and performance.

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2018-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are discussed. The authors demonstrated success in reaching desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and define a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2018-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nu- clear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and de ne a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

A simplified gross thrust computing technique for an afterburning turbofan engine

NASA Technical Reports Server (NTRS)

Hamer, M. J.; Kurtenbach, F. J.

1978-01-01

A simplified gross thrust computing technique extended to the F100-PW-100 afterburning turbofan engine is described. The technique uses measured total and static pressures in the engine tailpipe and ambient static pressure to compute gross thrust. Empirically evaluated calibration factors account for three-dimensional effects, the effects of friction and mass transfer, and the effects of simplifying assumptions for solving the equations. Instrumentation requirements and the sensitivity of computed thrust to transducer errors are presented. NASA altitude facility tests on F100 engines (computed thrust versus measured thrust) are presented, and calibration factors obtained on one engine are shown to be applicable to the second engine by comparing the computed gross thrust. It is concluded that this thrust method is potentially suitable for flight test application and engine maintenance on production engines with a minimum amount of instrumentation.

Castable thermal insulation for use as heat shields

NASA Technical Reports Server (NTRS)

Mountvala, A. J.; Nakamura, H. H.; Rechter, H. L.

1974-01-01

Structural members supporting the afterburners of high thrust rocket engines are subjected to extreme heating, along with severe vibration and high acceleration levels during early lift-off. Chemically-bonded, castable, zircon composite foams were developed and successfully tested to meet specific, laboratory simulated lift-off conditions.

RTE: A computer code for Rocket Thermal Evaluation

NASA Technical Reports Server (NTRS)

Naraghi, Mohammad H. N.

1995-01-01

The numerical model for a rocket thermal analysis code (RTE) is discussed. RTE is a comprehensive thermal analysis code for thermal analysis of regeneratively cooled rocket engines. The input to the code consists of the composition of fuel/oxidant mixture and flow rates, chamber pressure, coolant temperature and pressure. dimensions of the engine, materials and the number of nodes in different parts of the engine. The code allows for temperature variation in axial, radial and circumferential directions. By implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties. The fuel/oxidant mixture ratio can be varied along the thrust chamber. This feature allows the user to incorporate a non-equilibrium model or an energy release model for the hot-gas-side. The user has the option of bypassing the hot-gas-side calculations and directly inputting the gas-side fluxes. This feature is used to link RTE to a boundary layer module for the hot-gas-side heat flux calculations.

Evaluation of various thrust calculation techniques on an F404 engine

NASA Technical Reports Server (NTRS)

Ray, Ronald J.

1990-01-01

In support of performance testing of the X-29A aircraft at the NASA-Ames, various thrust calculation techniques were developed and evaluated for use on the F404-GE-400 engine. The engine was thrust calibrated at NASA-Lewis. Results from these tests were used to correct the manufacturer's in-flight thrust program to more accurately calculate thrust for the specific test engine. Data from these tests were also used to develop an independent, simplified thrust calculation technique for real-time thrust calculation. Comparisons were also made to thrust values predicted by the engine specification model. Results indicate uninstalled gross thrust accuracies on the order of 1 to 4 percent for the various in-flight thrust methods. The various thrust calculations are described and their usage, uncertainty, and measured accuracies are explained. In addition, the advantages of a real-time thrust algorithm for flight test use and the importance of an accurate thrust calculation to the aircraft performance analysis are described. Finally, actual data obtained from flight test are presented.

Numerical analysis of the hot-gas-side and coolant-side heat transfer in liquid rocket engine combustors

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Van, Luong

1992-01-01

The objective of this paper are to develop a multidisciplinary computational methodology to predict the hot-gas-side and coolant-side heat transfer and to use it in parametric studies to recommend optimized design of the coolant channels for a regeneratively cooled liquid rocket engine combustor. An integrated numerical model which incorporates CFD for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels, was developed. This integrated CFD/thermal model was validated by comparing predicted heat fluxes with those of hot-firing test and industrial design methods for a 40 k calorimeter thrust chamber and the Space Shuttle Main Engine Main Combustion Chamber. Parametric studies were performed for the Advanced Main Combustion Chamber to find a strategy for a proposed combustion chamber coolant channel design.

Space shuttle orbit maneuvering engine

NASA Technical Reports Server (NTRS)

Pauckert, R. P.

1975-01-01

Data on the performance, stability, and thermal characteristics of an OME operating with an alternate injector configuration and with alternate propellants was obtained. The design, manufacturing, and operating characteristics of an electroformed, regeneratively cooled thrust chamber were also derived. Subscale and full scale tests provide data relating to off-design and transient operation.

Evaluation of an Outer Loop Retrofit Architecture for Intelligent Turbofan Engine Thrust Control

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Sowers, T. Shane

2006-01-01

The thrust control capability of a retrofit architecture for intelligent turbofan engine control and diagnostics is evaluated. The focus of the study is on the portion of the hierarchical architecture that performs thrust estimation and outer loop thrust control. The inner loop controls fan speed so the outer loop automatically adjusts the engine's fan speed command to maintain thrust at the desired level, based on pilot input, even as the engine deteriorates with use. The thrust estimation accuracy is assessed under nominal and deteriorated conditions at multiple operating points, and the closed loop thrust control performance is studied, all in a complex real-time nonlinear turbofan engine simulation test bed. The estimation capability, thrust response, and robustness to uncertainty in the form of engine degradation are evaluated.

Architecture-Led Safety Process

DTIC Science & Technology

2016-12-01

Action Hazard Guide 42 Table 18: Comparative Table of Safety and Reliability Terms 47 CMU/SEI-2016-TR-012 | SOFTWARE ENGINEERING INSTITUTE...provides too much thrust Engine is slow to pro- vide commanded thrust (increase or de- crease) Engine will not shut- down when com - manded...Thrust level must be provided at the com - manded level H4: Engine is slow to provide commanded thrust SC3: Engine must provide commanded thrust in

A unique nuclear thermal rocket engine using a particle bed reactor

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Dahl, Wayne B.; McIlwain, Melvin C.

1992-01-01

Aerojet Propulsion Division (APD) studied 75-klb thrust Nuclear Thermal Rocket Engines (NTRE) with particle bed reactors (PBR) for application to NASA's manned Mars mission and prepared a conceptual design description of a unique engine that best satisfied mission-defined propulsion requirements and customer criteria. This paper describes the selection of a sprint-type Mars transfer mission and its impact on propulsion system design and operation. It shows how our NTRE concept was developed from this information. The resulting, unusual engine design is short, lightweight, and capable of high specific impulse operation, all factors that decrease Earth to orbit launch costs. Many unusual features of the NTRE are discussed, including nozzle area ratio variation and nozzle closure for closed loop after cooling. Mission performance calculations reveal that other well known engine options do not support this mission.

Extended performance solar electric propulsion thrust system study. Volume 4: Thruster technology evaluation

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Hawthorne, E. I.; Weisman, Y. C.; Frisman, M.; Benson, G. C.; Mcgrath, R. J.; Martinelli, R. M.; Linsenbardt, T. L.; Beattie, J. R.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. Emphasis was placed on relatively high power missions (60 to 100 kW) such as a Halley's comet rendezvous. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. The baseline thrust system design features modular construction, conventional power processing, and a concentrator solar array concept and is designed to interface with the Space Shuttle.

14 CFR 25.934 - Turbojet engine thrust reverser system tests.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbojet engine thrust reverser system... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant General § 25.934 Turbojet engine thrust reverser system tests. Thrust reversers installed on turbojet engines must meet the...

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2015-01-01

The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with maximizing the use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10-B2 engine and Delta Cryogenic Second Stage) to further ensure affordability. This paper provides a preliminary NASA, DOE and industry assessment of what is required - the key DDT&E activities, development options, and the associated schedule - to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 2: High pressure oxidizer turbo-pump turbine end bearing analysis

NASA Technical Reports Server (NTRS)

Sisk, Gregory A.

1989-01-01

The high-pressure oxidizer turbopump (HPOTP) consists of two centrifugal pumps, on a common shaft, that are directly driven by a hot-gas turbine. Pump shaft axial thrust is balanced in that the double-entry main inducer/impeller is inherently balanced and the thrusts of the preburner pump and turbine are nearly equal but opposite. Residual shaft thrust is controlled by a self-compensating, non-rubbing, balance piston. Shaft hang-up must be avoided if the balance piston is to perform properly. One potential cause of shaft hang-up is contact between the Phase 2 bearing support and axial spring cartridge of the HPOTP main pump housing. The status of the bearing support/axial spring cartridge interface is investigated under current loading conditions. An ANSYS version 4.3, three-dimensional, finite element model was generated on Lockheed's VAX 11/785 computer. A nonlinear thermal analysis was then executed on the Marshall Space Flight Center Engineering Analysis Data System (EADS). These thermal results were then applied along with the interference fit and bolt preloads to the model as load conditions for a static analysis to determine the gap status of the bearing support/axial spring cartridge interface. For possible further analysis of the local regions of HPOTP main pump housing assembly, detailed ANSYS submodels were generated using I-DEAS Geomod and Supertab (Appendix A).

Bimodal Nuclear Thermal Rocket Sizing and Trade Matrix for Lunar, Near Earth Asteroid and Mars Missions

NASA Astrophysics Data System (ADS)

McCurdy, David R.; Krivanek, Thomas M.; Roche, Joseph M.; Zinolabedini, Reza

2006-01-01

The concept of a human rated transport vehicle for various near earth missions is evaluated using a liquid hydrogen fueled Bimodal Nuclear Thermal Propulsion (BNTP) approach. In an effort to determine the preliminary sizing and optimal propulsion system configuration, as well as the key operating design points, an initial investigation into the main system level parameters was conducted. This assessment considered not only the performance variables but also the more subjective reliability, operability, and maintainability attributes. The SIZER preliminary sizing tool was used to facilitate rapid modeling of the trade studies, which included tank materials, propulsive versus an aero-capture trajectory, use of artificial gravity, reactor chamber operating pressure and temperature, fuel element scaling, engine thrust rating, engine thrust augmentation by adding oxygen to the flow in the nozzle for supersonic combustion, and the baseline turbopump configuration to address mission redundancy and safety requirements. A high level system perspective was maintained to avoid focusing solely on individual component optimization at the expense of system level performance, operability, and development cost.

Regeneratively cooled rocket engine for space storable propellants

NASA Technical Reports Server (NTRS)

Wagner, W. R.

1973-01-01

Analysis, design, fabrication, and test efforts were performed for the existing OF2/B2H6 regeneratively cooled lK (4448 N) thrust chamber to illustrate simultaneous B2H6 fuel and OF2 oxidizer cooling and to provide results for a gaseous propellant condition injected into the combustion chamber. Data derived from performance, thermal and flow measurements confirmed predictions derived from previous test work and from concurrent analytical study. Development data derived from the experimental study were indicated to be sufficient to develop a preflight thrust chamber demonstrator prototype for future space mission objectives.

14 CFR 23.934 - Turbojet and turbofan engine thrust reverser systems tests.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbojet and turbofan engine thrust... CATEGORY AIRPLANES Powerplant General § 23.934 Turbojet and turbofan engine thrust reverser systems tests. Thrust reverser systems of turbojet or turbofan engines must meet the requirements of § 33.97 of this...

Ion propulsion

NASA Technical Reports Server (NTRS)

Meserole, J. S.; Keefer, Dennis; Ruyten, Wilhelmus; Peng, Xiaohang

1995-01-01

An ion engine is a plasma thruster which produces thrust by extracting ions from the plasma and accelerating them to high velocity with an electrostatic field. The ions are then neutralized and leave the engine as high velocity neutral particles. The advantages of ion engines are high specific impulse and efficiency and their ability to operate over a wide range of input powers. In comparison with other electric thrusters, the ion engine has higher efficiency and specific impulse than thermal electric devices such as the arcjet, microwave, radiofrequency and laser heated thrusters and can operate at much lower current levels than the MPD thruster. However, the thrust level for an ion engine may be lower than a thermal electric thruster of the same operating power, consistent with its higher specific impulse, and therefore ion engines are best suited for missions which can tolerate longer duration propulsive phases. The critical issue for the ion engine is lifetime, since the prospective missions may require operation for several thousands of hours. The critical components of the ion engine, with respect to engine lifetime, are the screen and accelerating grid structures. Typically, these are large metal screens that must support a large voltage difference and maintain a small gap between them. Metallic whisker growth, distortion and vibration can lead to arcing, and over a long period of time ion sputtering will erode the grid structures and change their geometry. In order to study the effects of long time operation of the grid structure, we are developing computer codes based on the Particle-In-Cell (PIC) technique and Laser Induced Fluorescence (LIF) diagnostic techniques to study the physical processes which control the performance and lifetime of the grid structures.

Oxygen/Alcohol Dual Thrust RCS Engines

NASA Technical Reports Server (NTRS)

Angstadt, Tara; Hurlbert, Eric

1999-01-01

A non-toxic dual thrust RCS engine offers significant operational, safety, and performance advantages to the space shuttle and the next generation RLVs. In this concept, a single engine produces two thrust levels of 25 and 870 lbf. The low thrust level is provided by the spark torch igniter, which, with the addition of 2 extra valves, can also be made to function as a vernier. A dual thrust RCS engine allows 38 verniers to be packaged more efficiently on a vehicle. These 38 vemiers improve translation and reduce cross coupling, thereby providing more pure roll, pitch, and yaw maneuvers of the vehicle. Compared to the 6 vemiers currently on the shuttle, the 38 dual thrust engines would be 25 to 40% more efficient for the same maneuvers and attitude control. The vernier thrust level also reduces plume impingement and contamination concerns. Redundancy is also improved, thereby improving mission success reliability. Oxygen and ethanol are benign propellants which do not create explosive reaction products or contamination, as compared to hypergolic propellants. These characteristics make dual-thrust engines simpler to implement on a non-toxic reaction control system. Tests at WSTF in August 1999 demonstrated a dual-thrust concept that is successful with oxygen and ethanol. Over a variety of inlet pressures and mixture ratios at 22:1 area ratio, the engine produced between 230 and 297 sec Isp, and thrust levels from 8 lbf. to 50 lbf. This paper describes the benefits of dual-thrust engines and the recent results from tests at WSTF.

Rotational reflectance of dispersed vitrinite from the Arkoma basin

USGS Publications Warehouse

Houseknecht, D.W.; Weesner, C.M.B.

1997-01-01

Rotational reflectance of dispersed vitrinite provides superior documentation of thermal maturity and a capability for interpreting relative timing between thermal and kinematic events in Arkoma Basin strata characterized by vitrinite reflectances up to 5%. Rotational reflectance (R(rot)) is a more precise and less ambiguous index of thermal maturity than maximum (R'(max)), minimum (R(min)), and random (R(ran)) reflectance. Vitrinite reflectance anisotropy becomes sufficiently large to be measurable (using a microscope equipped with an automated rotating polarizer) at ???2% R(rot) and increases following a power function with increasing thermal maturity. Rotational reflectance data can be used to infer the shape of the vitrinite reflectance indicating surface (i.e. indicatrix) and, in turn, to enhance interpretations of the timing between thermal maxima and compressional tectonic events. Data from three wells in the Arkoma Basin Ouachita frontal thrust belt are used as examples. The absence of offsets in measured R(rot) across thrust faults combined with a predominance of uniaxial vitrinite in the thrust faulted part of the section suggest thermal maximum postdated thrust faulting in the western Ouachita frontal thrust belt of Oklahoma. In contrast, the general absence of offsets in measured R(rot) across thrust faults combined with a predominance of biaxial vitrinite in the thrust faulted part of the section suggest that the thermal maximum was coeval with thrust faulting in the eastern Ouachita frontal thrust belt of Arkansas. The presence of biaxial vitrinite in an allochthonous section and uniaxial vitrinite in an underlying, autochthonous section suggests that the thermal maximum was coeval with listric thrust faulting in the central Arkoma Basin of Oklahoma, and that rotational reflectance data can be used as a strain indicator to detect subtle decollement zones.

Experimental study on the thrust modulation performance of powdered magnesium and CO2 bipropellant engine

NASA Astrophysics Data System (ADS)

Li, Chao; Hu, Chunbo; Zhu, Xiaofei; Hu, Jiaming; Li, Yue; Hu, Xu

2018-06-01

Powdered Mg and CO2 bipropellant engine providing a practical demonstration of in situ resource utilization (ISRU) for Mars Sample Return (MSR) mission seems to be feasible by current investigations. However, essential functions of the engine to satisfy the complicated ballistics requirements such as thrust modulation and multiple pulse have not been established yet. The aim of this experimental study is to evaluate the engine's thrust modulation feasibility and to investigate its thrust modulation characteristics. A powdered Mg and CO2 bipropellant engine construction aiming to achieve thrust modulation ability was proposed. A mass flow rate calibration experiment to evaluate the gas-solid mass flow rate regulating performance was conducted before fire tests. Fire test result shows that the engine achieved successful ignition as well as self-sustaining combustion; Thrust modulation of the engine is feasible, detail thrust estimating result of the test shows that maximum thrust is 135.91 N and the minimum is 5.65 N with a 22.11 thrust modulation ratio, moreover, the transportation period is quick and the thrust modulation ratio is adjustable. At the same time, the powder feed system reaches a two-step flow rate regulating with a modulation ratio of 4.5-5. What' more, caused by the uneven engine working conditions, there is an obvious difference in combustion efficiency value, maximum combustion efficiency of the powdered Mg and CO2 bipropellant engine is 80.20%.

Magnetohydrodynamic Augmented Propulsion Experiment: I. Performance Analysis and Design

NASA Technical Reports Server (NTRS)

Litchford, R. J.; Cole, J. W.; Lineberry, J. T.; Chapman, J. N.; Schmidt, H. J.; Lineberry, C. W.

2003-01-01

The performance of conventional thermal propulsion systems is fundamentally constrained by the specific energy limitations associated with chemical fuels and the thermal limits of available materials. Electromagnetic thrust augmentation represents one intriguing possibility for improving the fuel composition of thermal propulsion systems, thereby increasing overall specific energy characteristics; however, realization of such a system requires an extremely high-energy-density electrical power source as well as an efficient plasma acceleration device. This Technical Publication describes the development of an experimental research facility for investigating the use of cross-field magnetohydrodynamic (MHD) accelerators as a possible thrust augmentation device for thermal propulsion systems. In this experiment,a 1.5-MW(sub e) Aerotherm arc heater is used to drive a 2-MW(sub e) MHD accelerator. The heatsink MHD accelerator is configured as an externally diagonalized, segmented channel, which is inserted into a large-bore, 2-T electromagnet. The performance analysis and engineering design of the flow path are described as well as the parameter measurements and flow diagnostics planned for the initial series of test runs.

Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets

NASA Technical Reports Server (NTRS)

Jones, William L.; Dowman, Harry W.

1947-01-01

Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Thermal design and analysis of a hydrogen-burning wind tunnel model of an airframe-integrated scramjet

NASA Technical Reports Server (NTRS)

Guy, R. W.; Mueller, J. N.; Pinckney, S. Z.; Lee, L. P.

1976-01-01

An aerodynamic model of a hydrogen burning, airframe integrated scramjet engine has been designed, fabricated, and instrumented. This model is to be tested in an electric arc heated wind tunnel at an altitude of 35.39 km (116,094 ft.) but with an inlet Mach number of 6 simulating precompression on an aircraft undersurface. The scramjet model is constructed from oxygen free, high conductivity copper and is a heat sink design except for water cooling in some critical locations. The model is instrumented for pressure, surface temperature, heat transfer rate, and thrust measurements. Calculated flow properties, heat transfer rates, and surface temperature distributions along the various engine components are included for the conditions stated above. For some components, estimates of thermal strain are presented which indicate significant reductions in plastic strain by selective cooling of the model. These results show that the 100 thermal cycle life of the engine was met with minimum distortion while staying within the 2669 N (600 lbf) engine weight limitation and while cooling the engine only in critical locations.

14 CFR 25.904 - Automatic takeoff thrust control system (ATTCS).

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Automatic takeoff thrust control system... Automatic takeoff thrust control system (ATTCS). Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when...

14 CFR 25.904 - Automatic takeoff thrust control system (ATTCS).

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Automatic takeoff thrust control system... Automatic takeoff thrust control system (ATTCS). Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when...

A Retro-Fit Control Architecture to Maintain Engine Performance With Usage

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Sowers, T. Shane; Garg, Sanjay

2007-01-01

An outer loop retrofit engine control architecture is presented which modifies fan speed command to obtain a desired thrust based on throttle position. This maintains the throttle-to-thrust relationship in the presence of engine degradation, which has the effect of changing the engine s thrust output for a given fan speed. Such an approach can minimize thrust asymmetry in multi-engine aircraft, and reduce pilot workload. The outer loop control is demonstrated under various levels of engine deterioration using a standard deterioration profile as well as an atypical profile. It is evaluated across various transients covering a wide operating range. The modified fan speed command still utilizes the standard engine control logic so all original life and operability limits remain in place. In all cases it is shown that with the outer loop thrust control in place, the deteriorated engine is able to match the thrust performance of a new engine up to the limits the controller will allow.

Aircraft Engine Thrust Estimator Design Based on GSA-LSSVM

NASA Astrophysics Data System (ADS)

Sheng, Hanlin; Zhang, Tianhong

2017-08-01

In view of the necessity of highly precise and reliable thrust estimator to achieve direct thrust control of aircraft engine, based on support vector regression (SVR), as well as least square support vector machine (LSSVM) and a new optimization algorithm - gravitational search algorithm (GSA), by performing integrated modelling and parameter optimization, a GSA-LSSVM-based thrust estimator design solution is proposed. The results show that compared to particle swarm optimization (PSO) algorithm, GSA can find unknown optimization parameter better and enables the model developed with better prediction and generalization ability. The model can better predict aircraft engine thrust and thus fulfills the need of direct thrust control of aircraft engine.

High-Heat-Flux Cyclic Durability of Thermal and Environmental Barrier Coatings

NASA Technical Reports Server (NTRS)

Zhu, Dongming; Ghosn, Louis L.; Miller, Robert A.

2007-01-01

Advanced ceramic thermal and environmental barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to protect the engine components and further raise engine temperatures. For the supersonic vehicles currently envisioned in the NASA fundamental aeronautics program, advanced gas turbine engines will be used to provide high power density thrust during the extended supersonic flight of the aircraft, while meeting stringent low emission requirements. Advanced ceramic coating systems are critical to the performance, life and durability of the hot-section components of the engine systems. In this work, the laser and burner rig based high-heat-flux testing approaches were developed to investigate the coating cyclic response and failure mechanisms under simulated supersonic long-duration cruise mission. The accelerated coating cracking and delamination mechanism under the engine high-heat-flux, and extended supersonic cruise time conditions will be addressed. A coating life prediction framework may be realized by examining the crack initiation and propagation in conjunction with environmental degradation under high-heat-flux test conditions.

Thrust and Propulsive Efficiency from an Instructive Viewpoint

ERIC Educational Resources Information Center

Kaufman, Richard D.

2010-01-01

In a typical engineering or physics curriculum, the momentum equation is used for the determination of jet engine thrust. Even a simple thrust analysis requires a heavy emphasis on mathematics that can cause students and engineers to lose a physical perspective on thrust. This article provides for this physical understanding using only static…

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.

Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

2016-01-01

This paper describes the process development for fabricating a high thermal conductivity NARloy-Z-Diamond composite (NARloy-Z-D) combustion chamber liner for application in advanced rocket engines. The fabrication process is challenging and this paper presents some details of these challenges and approaches used to address them. Prior research conducted at NASA-MSFC and Penn State had shown that NARloy-Z-40%D composite material has significantly higher thermal conductivity than the state of the art NARloy-Z alloy. Furthermore, NARloy-Z-40 %D is much lighter than NARloy-Z. These attributes help to improve the performance of the advanced rocket engines. Increased thermal conductivity will directly translate into increased turbopump power, increased chamber pressure for improved thrust and specific impulse. Early work on NARloy-Z-D composites used the Field Assisted Sintering Technology (FAST, Ref. 1, 2) for fabricating discs. NARloy-Z-D composites containing 10, 20 and 40vol% of high thermal conductivity diamond powder were investigated. Thermal conductivity (TC) data. TC increased with increasing diamond content and showed 50% improvement over pure copper at 40vol% diamond. This composition was selected for fabricating the combustion chamber liner using the FAST technique.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1981-01-01

An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

Design Considerations for Space Transfer Vehicles Using Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

Emrich, William J.

1995-01-01

The economical deployment of satellites to high energy earth orbits is crucial to the ultimate success of this nations commerical space ventures and is highly desirable for deep space planetary missions requiring earth escape trajectories. Upper stage space transfer vehicles needed to accomplish this task should ideally be simple, robust, and highly efficient. In this regard, solar thermal propulsion is particularly well suited to those missions where high thrust is not a requirement. The Marshall Space Flight Center is , therefore, currently engaged in defining a transfer vehicle employing solar thermal propulsion capable of transferring a 1000 lb. payload from low Earth orbit (LEO) to a geostationary Earth orbit (GEO) using a Lockheed launch vehicle (LLV3) with three Castors and a large shroud. The current design uses liquid hydrogen as the propellant and employs two inflatable 16 x 24 feet eliptical off-axis parabolic solar collectors to focus sunlight onto a tungsten/rhenium windowless black body type absorber. The concentration factor on this design is projected to be approximately 1800:1 for the primary collector and 2.42:1 for the secondary collector for an overall concentration factor of nearly 4400:1. The engine, which is about twice as efficient as the best currently available chemical engines, produces two pounds of thrust with a specific impulse (Isp) of 860 sec. Transfer times to GEO are projected to be on the order of one month. The launch and deployed configurations of the solar thermal upper stage (STUS) are depicted.

A study of variable thrust, variable specific impulse trajectories for solar system exploration

NASA Astrophysics Data System (ADS)

Sakai, Tadashi

A study has been performed to determine the advantages and disadvantages of variable thrust and variable Isp (specific impulse) trajectories for solar system exploration. There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse; or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality. This research has been conducted to answer the following questions: (1) Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion? (2) If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory? To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys. Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines; but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.

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).

Design and analysis report for the RL10-2B breadboard low thrust engine

NASA Technical Reports Server (NTRS)

Brown, J. R.; Foust, R. R.; Galler, D. E.; Kanic, P. G.; Kmiec, T. D.; Limerick, C. D.; Peckham, R. J.; Swartwout, T.

1984-01-01

The breadboard low thrust RL10-2B engine is described. A summary of the analysis and design effort to define the multimode thrust concept applicable to the requirements for the upper stage vehicles is provided. Baseline requirements were established for operation of the RL10-2B engine under the following conditions: (1) tank head idle at low propellant tank pressures without vehicle propellant conditioning or settling thrust; (2) pumped idle at a ten percent thrust level for low G deployment and/or vehicle tank pressurization; and (3) full thrust (15,000 lb.). Several variations of the engine configuration were investigated and results of the analyses are included.

Orbital transfer rocket engine technology: Advanced engine study

NASA Technical Reports Server (NTRS)

Hayden, Warren R.

1992-01-01

An advanced LOX/LH2 engine study for the use of NASA and vehicle prime contractors in developing concepts for manned missions to the Moon, Mars, and Phobos is documented. Parametric design data was obtained at five engine thrusts from 7.5K lbf to 50K lbf. Also, a separate task evaluated engine throttling over a 20:1 range and operation at a mixture ratio of 12 plus or minus 1 versus the 6 plus or minus 1 nominal. Cost data was also generated for DDT&E, first unit production, and factors in other life cycle costs. The major limitation of the study was lack of contact with vehicle prime contractors to resolve the issues in vehicle/engine interfaces. The baseline Aerojet dual propellant expander cycle was shown capable of meeting all performance requirements with an expected long operational life due to the high thermal margins. The basic engine design readily accommodated the 20:1 throttling requirement and operation up to a mixture ratio of 10 without change. By using platinum for baffled injector construction the increased thermal margin allowed operation up to mixture ratio 13. An initial engine modeling with an Aerojet transient simulation code (named MLETS) indicates stable engine operation with the baseline control system. A throttle ratio of 4 to 5 seconds from 10 percent to 100 percent thrust is also predicted. Performance predictions are 483.1 sec at 7.5K lbf, 487.3 sec at 20K lbf, and 485.2 sec at 50K lbf with a mixture ratio of 6 and an area ratio of 1200. Engine envelopes varied from 120 in. length/53 in. exit diameter at 7.5K lbf to 305 in. length/136 in. exit diameter at 50 K lbf. Packaging will be an important consideration. Continued work is recommended to include more vehicle prime contractor/engine contractor joint assessment of the interface issues.

Reverse thrust performance of the QCSEE variable pitch turbofan engine

NASA Technical Reports Server (NTRS)

Samanich, N. E.; Reemsnyder, D. C.; Blodmer, H. E.

1980-01-01

Results of steady state reverse and forward to reverse thrust transient performance tests are presented. The original quiet, clean, short haul, experimental engine four segment variable fan nozzle was retested in reverse and compared with a continuous, 30 deg half angle conical exlet. Data indicated that the significantly more stable, higher pressure recovery flow with the fixed 30 deg exlet resulted in lower engine vibrations, lower fan blade stress, and approximately a 20 percent improvement in reverse thrust. Objective reverse thrust of 35 percent of takeoff thrust was reached. Thrust response of less than 1.5 sec was achieved for the approach and the takeoff to reverse thrust transients.

NERVA-Derived Nuclear Thermal Propulsion Dual Mode Operation

NASA Astrophysics Data System (ADS)

Zweig, Herbert R.; Hundal, Rolv

1994-07-01

Generation of electrical power using the nuclear heat source of a NERVA-derived nuclear thermal rocket engine is presented. A 111,200 N thrust engine defined in a study for NASA-LeRC in FY92 is the reference engine for a three-engine vehicle for which a 50 kWe capacity is required. Processes are described for energy extraction from the reactor and for converting the energy to electricity. The tie tubes which support the reactor fuel elements are the source of thermal energy. The study focuses on process systems using Stirling cycle energy conversion operating at 980 K and an alternate potassium-Rankine system operating at 1,140 K. Considerations are given of the effect of the power production on turbopump operation, ZrH moderator dissociation, creep strain in the tie tubes, hydrogen permeation through the containment materials, requirements for a backup battery system, and the effects of potential design changes on reactor size and criticality. Nuclear considerations include changing tie tube materials to TZM, changing the moderator to low vapor-pressure yttrium hydride, and changing the fuel form from graphite matrix to a carbon-carbide composite.

The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Hoffman, Joe D.

1994-01-01

An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

Thrust modeling for hypersonic engines

NASA Technical Reports Server (NTRS)

Riggins, D. W.; Mcclinton, C. R.

1995-01-01

Expressions for the thrust losses of a scramjet engine are developed in terms of irreversible entropy increases and the degree of incomplete combustion. A method is developed which allows the calculation of the lost vehicle thrust due to different loss mechanisms within a given flow-field. This analysis demonstrates clearly the trade-off between mixing enhancement and resultant increased flow losses in scramjet combustors. An engine effectiveness parameter is defined in terms of thrust loss. Exergy and the thrust-potential method are related and compared.

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

A review of nuclear thermal propulsion carbide fuel corrosion and key issues

NASA Technical Reports Server (NTRS)

Pelaccio, Dennis G.; El-Genk, Mohamed S.

1994-01-01

Corrosion (mass loss) of carbide nuclear fuels due to their exposure to hot hydrogen in nuclear thermal propulsion engine systems greatly impacts the performance, thrust-to-weight and life of such systems. This report provides an overview of key issues and processes associated with the corrosion of carbide materials. Additionally, past pertinent development reactor test observations, as well as related experimental work and analysis modeling efforts are reviewed. At the conclusion, recommendations are presented, which provide the foundation for future corrosion modeling and verification efforts.

Solar Thermal Upper Stage Liquid Hydrogen Pressure Control Testing and Analytical Modeling

NASA Technical Reports Server (NTRS)

Olsen, A. D.; Cady, E. C.; Jenkins, D. S.; Chandler, F. O.; Grayson, G. D.; Lopez, A.; Hastings, L. J.; Flachbart, R. H.; Pedersen, K. W.

2012-01-01

The demonstration of a unique liquid hydrogen (LH2) storage and feed system concept for solar thermal upper stage was cooperatively accomplished by a Boeing/NASA Marshall Space Flight Center team. The strategy was to balance thermodynamic venting with the engine thrusting timeline during a representative 30-day mission, thereby, assuring no vent losses. Using a 2 cubic m (71 cubic ft) LH2 tank, proof-of-concept testing consisted of an engineering checkout followed by a 30-day mission simulation. The data were used to anchor a combination of standard analyses and computational fluid dynamics (CFD) modeling. Dependence on orbital testing has been incrementally reduced as CFD codes, combined with standard modeling, continue to be challenged with test data such as this.

Thrust Performance Evaluation of a Turbofan Engine Based on Exergetic Approach and Thrust Management in Aircraft

NASA Astrophysics Data System (ADS)

Yalcin, Enver

2017-05-01

The environmental parameters such as temperature and air pressure which are changing depending on altitudes are effective on thrust and fuel consumption of aircraft engines. In flights with long routes, thrust management function in airplane information system has a structure that ensures altitude and performance management. This study focused on thrust changes throughout all flight were examined by taking into consideration their energy and exergy performances for fuel consumption of an aircraft engine used in flight with long route were taken as reference. The energetic and exergetic performance evaluations were made under the various altitude conditions. The thrust changes for different altitude conditions were obtained to be at 86.53 % in descending direction and at 142.58 % in ascending direction while the energy and exergy efficiency changes for the referenced engine were found to be at 80.77 % and 84.45 %, respectively. The results revealed here can be helpful to manage thrust and reduce fuel consumption, but engine performance will be in accordance with operation requirements.

The 7.5K lbf thrust engine preliminary design for Orbit Transfer Vehicle

NASA Technical Reports Server (NTRS)

Hayden, Warren R.; Sabiers, Ralph; Schneider, Judy

1994-01-01

This document summarizes the preliminary design of the Aerojet version of the Orbit Transfer Vehicle main engine. The concept of a 7500 lbf thrust LO2/GH2 engine using the dual expander cycle for optimum efficiency is validated through power balance and thermal calculations. The engine is capable of 10:1 throttling from a nominal 2000 psia to a 200 psia chamber pressure. Reservations are detailed on the feasibility of a tank head start, but the design incorporates low speed turbopumps to mitigate the problem. The mechanically separate high speed turbopumps use hydrostatic bearings to meet engine life requirements, and operate at sub-critical speed for better throttling ability. All components were successfully packaged in the restricted envelope set by the clearances for the extendible/retractable nozzle. Gimbal design uses an innovative primary and engine out gimbal system to meet the +/- 20 deg gimbal requirement. The hydrogen regenerator and LOX/GH2 heat exchanger uses the Aerojet platelet structures approach for a highly compact component design. The extendible/retractable nozzle assembly uses an electric motor driven jack-screw design and a one segment carbon-carbon or silicide coated columbium nozzle with an area ratio, when extended, of 1430:1. A reliability analysis and risk assessment concludes the report.

Affordable Development and Demonstration of a Small Nuclear Thermal Rocket (NTR) Engine and Stage: How Small Is Big Enough?

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2016-01-01

The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 specific impulse - a 100 percent increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's Advanced Exploration Systems (AES) program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the Lead Fuel option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During fiscal year (FY) 2014, a preliminary Design Development Test and Evaluation (DDT&E) plan and schedule for NTP development was outlined by the NASA Glenn Research Center (GRC), Department of Energy (DOE) and industry that involved significant system-level demonstration projects that included Ground Technology Demonstration (GTD) tests at the Nevada National Security Site (NNSS), followed by a Flight Technology Demonstration (FTD) mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 kilopound-force thrust class, were considered. Both engine options used GC fuel and a common fuel element (FE) design. The small approximately 7.5 kilopound-force criticality-limited engine produces approximately157 thermal megawatts and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 kilopound-force Small Nuclear Rocket Engine (SNRE), developed by Los Alamos National Laboratory (LANL) at the end of the Rover program, produces approximately 367 thermal megawatts and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35-inch (approximately 89-centimeters) -long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 thermal megawatts of power. To reduce the cost of the FTD mission, a simple one-burn lunar flyby mission was considered to reduce the liquid hydrogen (LH2) propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids (NEA), and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Quiet Clean Short-haul Experimental Engine (QCSEE) over-the-wing engine and control simulation results

NASA Technical Reports Server (NTRS)

1978-01-01

A hybrid-computer simulation of the over the wing turbofan engine was constructed to develop the dynamic design of the control. This engine and control system includes a full authority digital electronic control using compressor stator reset to achieve fast thrust response and a modified Kalman filter to correct for sensor failures. Fast thrust response for powered-lift operations and accurate, fast responding, steady state control of the engine is provided. Simulation results for throttle bursts from 62 to 100 percent takeoff thrust predict that the engine will accelerate from 62 to 95 percent takeoff thrust in one second.

A design study of a reaction control system for a V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

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

1983-01-01

Attention is given to a short takeoff vertical landing (STOVL) aircraft reaction control system (RCS) design study. The STOVL fighter/attack aircraft employs an existing turbofan engine, and its hover requirement places a premium on weight reduction, which eliminates prospective nonairbreathing RCSs. A simple engine compressor bleed RCS degrades overall performance to an unacceptable degree, and the supersonic requirement precludes the large volume alternatives of thermal or ejector thrust augmentation systems as well as the ducting of engine exhaust gases and the use of a dedicated turbojet. The only system which addressed performance criteria without requiring major engine modifications was a dedicated load compressor driven by an auxilliary power unit.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

NASA Fixed Wing Project Propulsion Research and Technology Development Activities to Reduce Thrust Specific Energy Consumption

NASA Technical Reports Server (NTRS)

Hathaway, Michael D.; Rosario, Ruben Del; Madavan, Nateri K.

2013-01-01

This paper presents an overview of the propulsion research and technology portfolio of NASA Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 percent relative to a baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030 to 2035 time frame. The research investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts.

NASA Fixed Wing Project Propulsion Research and Technology Development Activities to Reduce Thrust Specific Energy Consumption

NASA Technical Reports Server (NTRS)

Hathaway, Michael D.; DelRasario, Ruben; Madavan, Nateri K.

2013-01-01

This paper presents an overview of the propulsion research and technology portfolio of NASA Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 % relative to a baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030-2035 time frame. The research investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

A Study on Aircraft Engine Control Systems for Integrated Flight and Propulsion Control

NASA Astrophysics Data System (ADS)

Yamane, Hideaki; Matsunaga, Yasushi; Kusakawa, Takeshi; Yasui, Hisako

The Integrated Flight and Propulsion Control (IFPC) for a highly maneuverable aircraft and a fighter-class engine with pitch/yaw thrust vectoring is described. Of the two IFPC functions the aircraft maneuver control utilizes the thrust vectoring based on aerodynamic control surfaces/thrust vectoring control allocation specified by the Integrated Control Unit (ICU) of a FADEC (Full Authority Digital Electronic Control) system. On the other hand in the Performance Seeking Control (PSC) the ICU identifies engine's various characteristic changes, optimizes manipulated variables and finally adjusts engine control parameters in cooperation with the Engine Control Unit (ECU). It is shown by hardware-in-the-loop simulation that the thrust vectoring can enhance aircraft maneuverability/agility and that the PSC can improve engine performance parameters such as SFC (specific fuel consumption), thrust and gas temperature.

Lateral dampers for thrust bearings

NASA Technical Reports Server (NTRS)

Hibner, D. H.; Szafir, D. R.

1985-01-01

The development of lateral damping schemes for thrust bearings was examined, ranking their applicability to various engine classes, selecting the best concept for each engine class and performing an in-depth evaluation. Five major engine classes were considered: large transport, military, small general aviation, turboshaft, and non-manrated. Damper concepts developed for evaluation were: curved beam, constrained and unconstrained elastomer, hybrid boost bearing, hydraulic thrust piston, conical squeeze film, and rolling element thrust face.

Thrust stand evaluation of engine performance improvement algorithms in an F-15 airplane

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1992-01-01

An investigation is underway to determine the benefits of a new propulsion system optimization algorithm in an F-15 airplane. The performance seeking control (PSC) algorithm optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses an onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. As part of the PSC test program, the F-15 aircraft was operated on a horizontal thrust stand. Thrust was measured with highly accurate load cells. The measured thrust was compared to onboard model estimates and to results from posttest performance programs. Thrust changes using the various PSC modes were recorded. Those results were compared to benefits using the less complex highly integrated digital electronic control (HIDEC) algorithm. The PSC maximum thrust mode increased intermediate power thrust by 10 percent. The PSC engine model did very well at estimating measured thrust and closely followed the transients during optimization. Quantitative results from the evaluation of the algorithms and performance calculation models are included with emphasis on measured thrust results. The report presents a description of the PSC system and a discussion of factors affecting the accuracy of the thrust stand load measurements.

Passively Adaptive Inflatable Structure for the Shooting Star Experiment

NASA Technical Reports Server (NTRS)

Tinker, Michael L..

1998-01-01

An inflatable structural system is described for the Shooting Star Experiment that is a technology demonstrator flight for solar thermal propulsion. The inflatable structure is a pressurized assembly used in orbit to support a fresnel lens for focusing sunlight into a thermal storage engine. When the engine temperature reaches a preset level, the propellant is injected into the storage engine, absorbs heat from a heat exchanger, and is expanded through the nozzle to produce thrust. The inflatable structure is an adaptive system in that a regulator and relief valve are utilized to maintain pressure within design limits during the full range of orbital conditions. Further, the polyimide film material used for construction of the inflatable is highly nonlinear, with modulus varying as a function of frequency, temperature, and level of excitation. A series of tests is described for characterizing the structure in response to various operating conditions.

A simple dynamic engine model for use in a real-time aircraft simulation with thrust vectoring

NASA Technical Reports Server (NTRS)

Johnson, Steven A.

1990-01-01

A simple dynamic engine model was developed at the NASA Ames Research Center, Dryden Flight Research Facility, for use in thrust vectoring control law development and real-time aircraft simulation. The simple dynamic engine model of the F404-GE-400 engine (General Electric, Lynn, Massachusetts) operates within the aircraft simulator. It was developed using tabular data generated from a complete nonlinear dynamic engine model supplied by the manufacturer. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. Included is a description of a method to account for axial thrust loss resulting from thrust vectoring. In addition, the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000 ft altitude and at Mach 0.7, 35,000 ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.

Centaur engine gimbal friction characteristics under simulated thrust load

NASA Technical Reports Server (NTRS)

Askew, J. W.

1986-01-01

An investigation was performed to determine the friction characteristics of the engine gimbal system of the Centaur upper stage rocket. Because the Centaur requires low-gain autopilots in order to meet all stability requirements for some configurations, control performance (response to transients and limit-cycle amplitudes) depends highly on these friction characteristics. Forces required to rotate the Centaur engine gimbal system were measured under a simulated thrust load of 66,723 N (15,000 lb) and in an altitude/thermal environment. A series of tests was performed at three test conditions; ambient temperature and pressure, ambient temperature and vacuum, and cryogenic temperature and vacuum. Gimbal rotation was controlled, and tests were performed in which rotation amplitude and frequency were varied by using triangular and sinusoidal waveforms. Test data revealed an elastic characteristic of the gimbal, independent of the input signal, which was evident prior to true gimbal sliding. The torque required to initiate gimbal sliding was found to decrease when both pressure and temperature decreased. Results from the low amplitude and low frequency data are currently being used in mathematically modeling the gimbal friction characteristics for Centaur autopilot performance studies.

Centaur engine gimbal friction characteristics under simulated thrust load

NASA Astrophysics Data System (ADS)

Askew, J. W.

1986-09-01

An investigation was performed to determine the friction characteristics of the engine gimbal system of the Centaur upper stage rocket. Because the Centaur requires low-gain autopilots in order to meet all stability requirements for some configurations, control performance (response to transients and limit-cycle amplitudes) depends highly on these friction characteristics. Forces required to rotate the Centaur engine gimbal system were measured under a simulated thrust load of 66,723 N (15,000 lb) and in an altitude/thermal environment. A series of tests was performed at three test conditions; ambient temperature and pressure, ambient temperature and vacuum, and cryogenic temperature and vacuum. Gimbal rotation was controlled, and tests were performed in which rotation amplitude and frequency were varied by using triangular and sinusoidal waveforms. Test data revealed an elastic characteristic of the gimbal, independent of the input signal, which was evident prior to true gimbal sliding. The torque required to initiate gimbal sliding was found to decrease when both pressure and temperature decreased. Results from the low amplitude and low frequency data are currently being used in mathematically modeling the gimbal friction characteristics for Centaur autopilot performance studies.

Breadboard RL10-11B low thrust operating mode

NASA Technical Reports Server (NTRS)

Kmiec, Thomas D.; Galler, Donald E.

1987-01-01

Cryogenic space engines require a cooling process to condition engine hardware to operating temperature before start. This can be accomplished most efficiently by burning propellants that would otherwise be dumped overboard after cooling the engine. The resultant low thrust operating modes are called Tank Head Idle and Pumped Idle. During February 1984, Pratt & Whitney conducted a series of tests demonstrating operation of the RL10 rocket engines at low thrust levels using a previously untried hydrogen/oxygen heat exchanger. The initial testing of the RL10-11B Breadboard Low Thrust Engine is described. The testing demonstrated operation at both tank head idle and pumped idle modes.

Evaluation of a simplified gross thrust calculation method for a J85-21 afterburning turbojet engine in an altitude facility

NASA Technical Reports Server (NTRS)

Baer-Riedhart, J. L.

1982-01-01

A simplified gross thrust calculation method was evaluated on its ability to predict the gross thrust of a modified J85-21 engine. The method used tailpipe pressure data and ambient pressure data to predict the gross thrust. The method's algorithm is based on a one-dimensional analysis of the flow in the afterburner and nozzle. The test results showed that the method was notably accurate over the engine operating envelope using the altitude facility measured thrust for comparison. A summary of these results, the simplified gross thrust method and requirements, and the test techniques used are discussed in this paper.

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

NASA Technical Reports Server (NTRS)

Henneberry, Hugh M.; Snyder, Christopher A.

1993-01-01

An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

High-speed engine/component performance assessment using exergy and thrust-based methods

NASA Technical Reports Server (NTRS)

Riggins, D. W.

1996-01-01

This investigation summarizes a comparative study of two high-speed engine performance assessment techniques based on energy (available work) and thrust-potential (thrust availability). Simple flow-fields utilizing Rayleigh heat addition and one-dimensional flow with friction are used to demonstrate the fundamental inability of conventional energy techniques to predict engine component performance, aid in component design, or accurately assess flow losses. The use of the thrust-based method on these same examples demonstrates its ability to yield useful information in all these categories. Energy and thrust are related and discussed from the stand-point of their fundamental thermodynamic and fluid dynamic definitions in order to explain the differences in information obtained using the two methods. The conventional definition of energy is shown to include work which is inherently unavailable to an aerospace Brayton engine. An engine-based energy is then developed which accurately accounts for this inherently unavailable work; performance parameters based on this quantity are then shown to yield design and loss information equivalent to the thrust-based method.

The nuclear thermal electric rocket: a proposed innovative propulsion concept for manned interplanetary missions

NASA Astrophysics Data System (ADS)

Dujarric, C.; Santovincenzo, A.; Summerer, L.

2013-03-01

Conventional propulsion technology (chemical and electric) currently limits the possibilities for human space exploration to the neighborhood of the Earth. If farther destinations (such as Mars) are to be reached with humans on board, a more capable interplanetary transfer engine featuring high thrust, high specific impulse is required. The source of energy which could in principle best meet these engine requirements is nuclear thermal. However, the nuclear thermal rocket technology is not yet ready for flight application. The development of new materials which is necessary for the nuclear core will require further testing on ground of full-scale nuclear rocket engines. Such testing is a powerful inhibitor to the nuclear rocket development, as the risks of nuclear contamination of the environment cannot be entirely avoided with current concepts. Alongside already further matured activities in the field of space nuclear power sources for generating on-board power, a low level investigation on nuclear propulsion has been running since long within ESA, and innovative concepts have already been proposed at an IAF conference in 1999 [1, 2]. Following a slow maturation process, a new concept was defined which was submitted to a concurrent design exercise in ESTEC in 2007. Great care was taken in the selection of the design parameters to ensure that this quite innovative concept would in all respects likely be feasible with margins. However, a thorough feasibility demonstration will require a more detailed design including the selection of appropriate materials and the verification that these can withstand the expected mechanical, thermal, and chemical environment. So far, the predefinition work made clear that, based on conservative technology assumptions, a specific impulse of 920 s could be obtained with a thrust of 110 kN. Despite the heavy engine dry mass, a preliminary mission analysis using conservative assumptions showed that the concept was reducing the required Initial Mass in Low Earth Orbit compared to conventional nuclear thermal rockets for a human mission to Mars. Of course, the realization of this concept still requires proper engineering and the dimensioning of quite unconventional machinery. A patent was filed on the concept. Because of the operating parameters of the nuclear core, which are very specific to this type of concept, it seems possible to test on ground this kind of engine at full scale in close loop using a reasonable size test facility with safe and clean conditions. Such tests can be conducted within fully confined enclosure, which would substantially increase the associated inherent nuclear safety levels. This breakthrough removes a showstopper for nuclear rocket engines development. The present paper will disclose the NTER (Nuclear Thermal Electric Rocket) engine concept, will present some of the results of the ESTEC concurrent engineering exercise, and will explain the concept for the NTER on-ground testing facility. Regulations and safety issues related to the development and implementation of the NTER concept will be addressed as well.

Real-time in-flight thrust calculation on a digital electronic engine control-equipped F100 engine in an F-15 airplane

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1984-01-01

Computer algorithms which calculate in-flight engine and aircraft performance real-time are discussed. The first step was completed with the implementation of a real-time thrust calculation program on a digital electronic engine control (DEEC) equiped F100 engine in an F-15 aircraft. The in-flight thrust modifications that allow calculations to be performed in real-time, to compare results to predictions, are presented.

Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Anderson, William E.

2005-01-01

The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

Hydrogen-oxygen auxiliary propulsion for the space shuttle. Volume 1: High pressure thrusters

NASA Technical Reports Server (NTRS)

1973-01-01

Technology for long life, high performing, gaseous hydrogen-gaseous oxygen rocket engines suitable for auxiliary propulsion was provided by a combined analytical and experimental program. Propellant injectors, fast response valves, igniters, and regeneratively and film-cooled thrust chambers were tested over a wide range of operating conditions. Data generated include performance, combustion efficiency, thermal characteristics film cooling effectiveness, dynamic response in pulsing, and cycle life limitations.

A design perspective on thermal barrier coatings

NASA Astrophysics Data System (ADS)

Soechting, F. O.

1999-12-01

This article addresses the challenges for maximizing the benefit of thermal barrier coatings for turbine engine applications. The perspective is from the viewpoint of a customer, a turbine airfoil designer who is continuously challenged to increase the turbine inlet temperature capability for new products while maintaining cooling flow levels or even reducing them. This is a fundamental requirement for achieving increased engine thrust levels. Developing advanced material systems for the turbine flowpath airfoils, such as high-temperature nickel-base superalloys or thermal barrier coatings to insulate the metal airfoils from the hot flowpath environment, is one approach to solve this challenge. The second approach is to increase the cooling performance of the turbine airfoil, which enables increased flowpath temperatures and reduced cooling flow levels. Thermal barrier coatings have been employed in jet engine applications for almost 30 years. The initial application was on augmentor liners to provide thermal protection during afterburner operation. However, the production use of thermal barrier coatings in the turbine section has only occurred in the past 15 years. The application was limited to stationary parts and only recently incorporated on the rotating turbine blades. This lack of endorsement of thermal barrier coatings resulted from the poor initial duratbility of these coatings in high heat flux environments. Significant improvements have been made to enhance spallation resistance and erosion resistance, which has resulted in increased reliability of these coatings in turbine applications.

The NASA Advanced Exploration Systems Nuclear Thermal Propulsion Project

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Mitchell, Doyce P.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Gerrish, Harold P.; Doughty, Glen; Belvin, Anthony; Clement, Steven; Borowski, Stanley K.;

Thrust Area Report, Engineering Research, Development and Technology

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

Langland, R. T.

1997-02-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the knowledge base, process technologies, specialized equipment, tools and facilities to support current and future LLNL programs. Engineering`s efforts are guided by a strategy that results in dual benefit: first, in support of Department of Energy missions, such as national security through nuclear deterrence; and second, in enhancing the nation`s economic competitiveness through our collaboration with U.S. industry in pursuit of the most cost- effective engineering solutions to LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Programmore » has two important goals: (1) identify key technologies relevant to LLNL programs where we can establish unique competencies, and (2) conduct high-quality research and development to enhance our capabilities and establish ourselves as the world leaders in these technologies. To focus Engineering`s efforts technology {ital thrust areas} are identified and technical leaders are selected for each area. The thrust areas are comprised of integrated engineering activities, staffed by personnel from the nine electronics and mechanical engineering divisions, and from other LLNL organizations. This annual report, organized by thrust area, describes Engineering`s activities for fiscal year 1996. The report provides timely summaries of objectives, methods, and key results from eight thrust areas: Computational Electronics and Electromagnetics; Computational Mechanics; Microtechnology; Manufacturing Technology; Materials Science and Engineering; Power Conversion Technologies; Nondestructive Evaluation; and Information Engineering. Readers desiring more information are encouraged to contact the individual thrust area leaders or authors. 198 refs., 206 figs., 16 tabs.« less

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.

Measuring Model Rocket Engine Thrust Curves

ERIC Educational Resources Information Center

Penn, Kim; Slaton, William V.

2010-01-01

This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

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

NASA Technical Reports Server (NTRS)

Kubiak, Jonathan M.; Arnett, Lori A.

2016-01-01

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

Conceptual Design of a Z-Pinch Fusion Propulsion System

NASA Technical Reports Server (NTRS)

Adams, Robert; Polsgrove, Tara; Fincher, Sharon; Fabinski, Leo; Maples, Charlotte; Miernik, Janie; Stratham, Geoffrey; Cassibry, Jason; Cortez, Ross; Turner, Matthew;

High Pressure Earth Storable Rocket Technology Program: Basic Program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Sicher, D.; Huang, D.; Mueller, T.

1995-01-01

The HIPES Program was conducted for NASA-LeRC by TRW. The Basic Program consisted of system studies, design of testbed engine, fabrication and testing of engine. Studies of both pressure-fed and pump-fed systems were investigated for N2O4 and both MMH and N2H4 fuels with the result that N2H4 provides the maximum payload for all satellites over MMH. The higher pressure engine offers improved performance with smaller envelope and associated weight savings. Pump-fed systems offer maximum payload for large and medium weight satellites while pressure-fed systems offer maximum payload for small light weight satellites. The major benefits of HIPES are high performance within a confined length maximizing payload for lightsats which are length (volume) constrained. Three types of thrust chambers were evaluated -- Copper heatsink at 400, 500 and 600 psia chamber pressures for performance/thermal; water cooled to determine heat absorbed to predict rhenium engine operation; and rhenium to validate the concept. The HIPES engine demonstrated very high performance at 50 lbf thrust (epsilon = 150) and Pc = 500 psia with both fuels: Isp = 337 sec using N2O4-N2H4 and ISP = 327.5 sec using N2O4-MMH indicating combustion efficiencies greater than 98%. A powder metallurgy rhenium engine demonstrated operation with high performance at Pc = 500 psia which indicated the viability of the concept.

Present capabilities and future requirements for computer-aided geometric modeling in the design and manufacture of gas turbine

NASA Technical Reports Server (NTRS)

Caille, E.; Propen, M.; Hoffman, A.

1984-01-01

Gas turbine engine design requires the ability to rapidly develop complex structures which are subject to severe thermal and mechanical operating loads. As in all facets of the aerospace industry, engine designs are constantly driving towards increased performance, higher temperatures, higher speeds, and lower weight. The ability to address such requirements in a relatively short time frame has resulted in a major thrust towards integrated design/analysis/manufacturing systems. These computer driven graphics systems represent a unique challenge, with major payback opportunities if properly conceived, implemented, and applied.

Parametric (On-Design) Cycle Analysis for a Separate-Exhaust Turbofan Engine With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Siow, Y. K.; Marek, C. J.

2005-01-01

Today s modern aircraft is based on air-breathing jet propulsion systems, which use moving fluids as substances to transform energy carried by the fluids into power. Throughout aero-vehicle evolution, improvements have been made to the engine efficiency and pollutants reduction. The major advantages associated with the addition of ITB are an increase in thermal efficiency and reduction in NOx emission. Lower temperature peak in the main combustor results in lower thermal NOx emission and lower amount of cooling air required. This study focuses on a parametric (on-design) cycle analysis of a dual-spool, separate-flow turbofan engine with an Interstage Turbine Burner (ITB). The ITB considered in this paper is a relatively new concept in modern jet engine propulsion. The ITB serves as a secondary combustor and is located between the high- and the low-pressure turbine, i.e., the transition duct. The objective of this study is to use design parameters, such as flight Mach number, compressor pressure ratio, fan pressure ratio, fan bypass ratio, and high-pressure turbine inlet temperature to obtain engine performance parameters, such as specific thrust and thrust specific fuel consumption. Results of this study can provide guidance in identifying the performance characteristics of various engine components, which can then be used to develop, analyze, integrate, and optimize the system performance of turbofan engines with an ITB. Visual Basic program, Microsoft Excel macrocode, and Microsoft Excel neuron code are used to facilitate Microsoft Excel software to plot engine performance versus engine design parameters. This program computes and plots the data sequentially without forcing users to open other types of plotting programs. A user s manual on how to use the program is also included in this report. Furthermore, this stand-alone program is written in conjunction with an off-design program which is an extension of this study. The computed result of a selected design-point engine will be exported to an engine reference data file that is required in off-design calculation.

Effects of bleed air extraction on thrust levels on the F404-GE-400 turbofan engine

NASA Technical Reports Server (NTRS)

Yuhas, Andrew J.; Ray, Ronald J.

1992-01-01

A ground test was performed to determine the effects of compressor bleed flow extraction on the performance of F404-GE-400 afterburning turbofan engines. The two engines were installed in the F/A-18 High Alpha Research Vehicle at the NASA Dryden Flight Research Facility. A specialized bleed ducting system was installed onto the aircraft to control and measure engine bleed airflow while the aircraft was tied down to a thrust measuring stand. The test was conducted on each engine and at various power settings. The bleed air extraction levels analyzed included flow rates above the manufacturer's maximum specification limit. The measured relationship between thrust and bleed flow extraction was shown to be essentially linear at all power settings with an increase in bleed flow causing a corresponding decrease in thrust. A comparison with the F404-GE-400 steady-state engine simulation showed the estimation to be within +/- 1 percent of measured thrust losses for large increases in bleed flow rate.

Short-term performance deterioration in JT9D-7A(SP) engine 695743

NASA Technical Reports Server (NTRS)

Bouchard, R. J.; Beyerly, W. R.; Sallee, G. P.

1978-01-01

The level of performance deterioration that occurred during early service was studied and also the contribution of each engine module to the overall level of engine performance deterioration. To assist in the distribution of the performance losses to the module level, testing with expanded experimental instrumentation was accomplished to permit direct measurement of module performance. An analytical teardown of the engine was accomplished to permit assignment of module performance losses to individual mechanical damage mechanisms including blade and seal wear, roughness, and thermal distortion of turbine parts. The results of the testing show that the engine deteriorated 1.5 percent in thrust specific fuel consumption from its performance when new. Initial testing, which included water washing, showed that 0.2 percent in performance deterioration was due to surface contamination (dirt) and 0.1 percent was due to drift of the engine vane control system, leaving 1.2 percent in basic TSFC deterioration. This 1.2 percent TSFC loss was distributed among the engine modules with 0.6 percent caused by clearance changes, 0.4 percent loss due to thermal distortion, and 0.2 percent loss due to increased roughness of the fan and low-pressure compressor airfoils.

On-Board Chemical Propulsion

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1997-01-01

NASA Lewis Research Center's On-Board Propulsion program (OBP) is developing low-thrust chemical propulsion technologies for both satellite and vehicle reaction control applications. There is a vigorous international competition to develop new, highperformance bipropellant engines. High-leverage bipropellant systems are critical to both commercial competitiveness in the international communications market and to cost-effective mission design in government sectors. To significantly improve bipropellant engine performance, we must increase the thermal margin of the chamber materials. Iridium-coated rhenium (Ir/Re) engines, developed and demonstrated under OBP programs, can operate at temperatures well above the constraints of state-of-practice systems, providing a sufficient margin to maximize performance with the hypergolic propellants used in most satellite propulsion systems.

Parameter-Study Of The Thermal Yarkovsky Effect Acting On Neas

NASA Astrophysics Data System (ADS)

Polishook, David; Prialnik, D.; Rosenberg, E.; Brosch, N.

2010-10-01

We study the relevant parameters for the thermal Yarkovsky effect acting on Near-Earth Asteroids (NEAs), specifically the rotation period and rotation axis. The study uses a quasi 3-D thermal model to derive the temperature map over the surface of the asteroid, and the thermal thrust is calculated. The model (Prialnik et al. 2004, Rosenberg and Prialnik 2006), uses an implicit scheme to numerically solve the equations that describe the asteroid and its thermal evolution. The results show how the thermal thrust is stronger for fast-rotating asteroids, as heat is emitted from their surface on the evening side, increasing the tangential component of the thermal thrust. Moreover, we show the differences in the thermal thrust between asteroids with different perihelion distances, and how this can explain the observed distribution of asteroids in the inner Solar System on the spin-perihelion plane. Our results suggest that many asteroids within the inner Solar System may have retrograde spins. Acknowledgements: D. Polishook is grateful for an Ilan Ramon doctoral scholarship from the Israeli Ministry of Science.

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

NASA Technical Reports Server (NTRS)

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

1982-01-01

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

Hamilton Standard Q-fan demonstrator dynamic pitch change test program, volume 1

NASA Technical Reports Server (NTRS)

Demers, W. J.; Nelson, D. J.; Wainauski, H. S.

1975-01-01

Tests of a full scale variable pitch fan engine to obtain data on the structural characteristics, response times, and fan/core engine compatibility during transient changes in blade angle, fan rpm, and engine power is reported. Steady state reverse thrust tests with a take off nozzle configuration were also conducted. The 1.4 meter diameter, 13 bladed controllable pitch fan was driven by a T55 L 11A engine with power and blade angle coordinated by a digital computer. The tests demonstrated an ability to change from full forward thrust to reverse thrust in less than one (1) second. Reverse thrust was effected through feather and through flat pitch; structural characteristics and engine/fan compatibility were within satisfactory limits.

Synthesis, Characterization And Modeling Of Functionally Graded Multifunctional Hybrid Composites For Extreme Environments

DTIC Science & Technology

2017-04-04

research thrust areas are designed to enable the development of reliable, damage tolerant, lightweight structures with excellent thermal management...46 2. RESEARCH THRUST AREA: MULTISCALE CHARACTERIZATION AND MODELING .................................... 56 2.1 DESIGN OF MATERIALS...The research thrust areas are designed to enable the development of reliable, damage tolerant, lightweight structures with excellent thermal

Propulsion system needs

NASA Technical Reports Server (NTRS)

Gunn, Stanley

1991-01-01

The needs of the designer of a solid core nuclear rocket engine are discussed. Some of the topics covered include: (1) a flight thrust module/feed system module assembly; (2) a nuclear thermal rocket (NTR), expander cycle, dual T/P; (3) turbopump operating conditions; (4) typical system parameters; (5) growth capability composite fuel elements; (6) a NTR radiation cooled nozzle extension; (7) a NFS-3B Feed System; and (8) a NTR Integrated Pneumatic-Fluidics Control System.

Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough?

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg (Abraham); Joyner, Claude R.

2015-01-01

The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 seconds - a 100% increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's AES program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the "Lead Fuel" option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During FY'14, a preliminary DDT&E plan and schedule for NTP development was outlined by GRC, DOE and industry that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 klbf thrust class, were considered. Both engine options used GC fuel and a "common" fuel element (FE) design. The small approximately 7.5 klbf "criticality-limited" engine produces approximately 157 megawatts of thermal power (MWt) and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 klbf Small Nuclear Rocket Engine (SNRE), developed by LANL at the end of the Rover program, produces approximately 367 MWt and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35 inch (approximately 89 cm) long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 MWt of power. To reduce the cost of the FTD mission, a simple "1-burn" lunar flyby mission was considered to reduce the LH2 propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids, and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Thrust chamber material technology program

NASA Technical Reports Server (NTRS)

Andrus, J. S.; Bordeau, R. G.

1989-01-01

This report covers work performed at Pratt & Whitney on development of copper-based materials for long-life, reusable, regeneratively cooled rocket engine thrust chambers. The program approached the goal of enhanced cyclic life through the application of rapid solidification to alloy development, to introduce fine dispersions to strengthen and stabilize the alloys at elevated temperatures. After screening of alloy systems, copper-based alloys containing Cr, Co, Hf, Ag, Ti, and Zr were processed by rapid-solidification atomization in bulk quantities. Those bulk alloys showing the most promise were characterized by tensile testing, thermal conductivity testing, and elevated-temperature, low-cycle fatigue (LFC) testing. Characterization indicated that Cu- 1.1 percent Hf exhibited the greatest potential as an improved-life thrust chamber material, exhibiting LCF life about four times that of NASA-Z. Other alloys (Cu- 0.6 percent Zr, and Cu- 0.6 percent Zr- 1.0 percent Cr) exhibited promise for use in this application, but needed more development work to balance properties.

Design of the laser acupuncture therapeutic instrument.

PubMed

Li, Chengwei; Zhen, Huang

2006-01-01

Laser acupuncture is defined as the stimulation of traditional acupuncture points with low-intensity, non-thermal laser irradiation. It has been well applied in clinic since the 1970s; however, some traditional acupuncture manipulating methods still cannot be implemented in the design of this kind of instruments, such as lifting and thrusting manipulating method, and twisting and twirling manipulating method, which are the essential acupuncture method in traditional acupuncture. The objective of this work was to design and build a low cost portable laser acupuncture therapeutic instrument, which can implement the two essential acupuncture manipulating methods. Digital PID control theory is used to control the power of laser diode (LD), and to implement the lifting and thrusting manipulating method. Special optical system is designed to implement twisting and twirling manipulating method. M5P430 microcontroller system is used as the control centre of the instrument. The realization of lifting and thrusting manipulating method and twisting and twirling manipulating method are technological innovations in traditional acupuncture coming true in engineering.

Tests of a D vented thrust deflecting nozzle behind a simulated turbofan engine

NASA Technical Reports Server (NTRS)

Watson, T. L.

1982-01-01

A D vented thrust deflecting nozzle applicable to subsonic V/STOL aircraft was tested behind a simulated turbofan engine in the verticle thrust stand. Nozzle thrust, fan operating characteristics, nozzle entrance conditions, and static pressures were measured. Nozzle performance was measured for variations in exit area and thrust deflection angle. Six core nozzle configurations, the effect of core exit axial location, mismatched core and fan stream nozzle pressure ratios, and yaw vane presence were evaluated. Core nozzle configuration affected performance at normal and engine out operating conditions. Highest vectored nozzle performance resulted for a given exit area when core and fan stream pressure were equal. Its is concluded that high nozzle performance can be maintained at both normal and engine out conditions through control of the nozzle entrance Mach number with a variable exit area.

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

Reusable thrust chamber and injector concepts were evaluated for the space shuttle orbit maneuvering engine (OME). Parametric engine calculations were carried out by computer program for N2O4/amine, LOX/amine and LOX/hydrocarbon propellant combinations for engines incorporating regenerative cooled and insulated columbium thrust chambers. The calculation methods are described including the fuel vortex film cooling method of combustion gas temperature control, and performance prediction. A method of acceptance of a regeneratively cooled heat rejection reduction using a silicone oil additive was also demonstrated by heated tube heat transfer testing. Regeneratively cooled thrust chamber operation was also demonstrated where the injector was characterized for the OME application with a channel wall regenerative thrust chamber. Bomb stability testing of the demonstration chambers/injectors demonstrated recovery for the nominal design of acoustic cavities. Cavity geometry changes were also evaluated to assess their damping margin. Performance and combustion stability was demonstrated of the originally developed 10 inch diameter combustion pattern operating in an 8 inch diameter thrust chamber.

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).

Characterization of lubrication oil emissions from aircraft engines.

PubMed

Yu, Zhenhong; Liscinsky, David S; Winstead, Edward L; True, Bruce S; Timko, Michael T; Bhargava, Anuj; Herndon, Scott C; Miake-Lye, Richard C; Anderson, Bruce E

2010-12-15

In this first ever study, particulate matter (PM) emitted from the lubrication system overboard breather vent for two different models of aircraft engines has been systematically characterized. Lubrication oil was confirmed as the predominant component of the emitted particulate matter based upon the characteristic mass spectrum of the pure oil. Total particulate mass and size distributions of the emitted oil are also investigated by several high-sensitivity aerosol characterization instruments. The emission index (EI) of lubrication oil at engine idle is in the range of 2-12 mg kg(-1) and increases with engine power. The chemical composition of the oil droplets is essentially independent of engine thrust, suggesting that engine oil does not undergo thermally driven chemical transformations during the ∼4 h test window. Volumetric mean diameter is around 250-350 nm for all engine power conditions with a slight power dependence.

Full Flight Envelope Direct Thrust Measurement on a Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Conners, Timothy R.; Sims, Robert L.

1998-01-01

Direct thrust measurement using strain gages offers advantages over analytically-based thrust calculation methods. For flight test applications, the direct measurement method typically uses a simpler sensor arrangement and minimal data processing compared to analytical techniques, which normally require costly engine modeling and multisensor arrangements throughout the engine. Conversely, direct thrust measurement has historically produced less than desirable accuracy because of difficulty in mounting and calibrating the strain gages and the inability to account for secondary forces that influence the thrust reading at the engine mounts. Consequently, the strain-gage technique has normally been used for simple engine arrangements and primarily in the subsonic speed range. This paper presents the results of a strain gage-based direct thrust-measurement technique developed by the NASA Dryden Flight Research Center and successfully applied to the full flight envelope of an F-15 aircraft powered by two F100-PW-229 turbofan engines. Measurements have been obtained at quasi-steady-state operating conditions at maximum non-augmented and maximum augmented power throughout the altitude range of the vehicle and to a maximum speed of Mach 2.0 and are compared against results from two analytically-based thrust calculation methods. The strain-gage installation and calibration processes are also described.

Statistical error model for a solar electric propulsion thrust subsystem

NASA Technical Reports Server (NTRS)

Bantell, M. H.

1973-01-01

The solar electric propulsion thrust subsystem statistical error model was developed as a tool for investigating the effects of thrust subsystem parameter uncertainties on navigation accuracy. The model is currently being used to evaluate the impact of electric engine parameter uncertainties on navigation system performance for a baseline mission to Encke's Comet in the 1980s. The data given represent the next generation in statistical error modeling for low-thrust applications. Principal improvements include the representation of thrust uncertainties and random process modeling in terms of random parametric variations in the thrust vector process for a multi-engine configuration.

Computation of optimal Mars trajectories via combined chemical/electrical propulsion, Part 3: Compromise solutions

NASA Astrophysics Data System (ADS)

Miele, A.; Wang, T.; Williams, P. N.

2005-12-01

The success of the solar-electric ion engine powering the DS1 spacecraft has paved the way toward the use of low-thrust electrical engines in future planetary/interplanetary missions. Vis-à-vis a chemical engine, an electrical engine has a higher specific impulse, implying a possible decrease in propellant mass; however, the low-thrust aspect discourages the use of an electrical engine in the near-planet phases of a trip, since this might result in an increase in flight time. Therefore, a fundamental design problem is to find the best combination of chemical propulsion and electrical propulsion for a given mission, for example, a mission from Earth to Mars. With this in mind, this paper is the third of a series dealing with the optimization of Earth Mars missions via the use of hybrid engines, namely the combination of high-thrust chemical engines for planetary flight and low-thrust electrical engines for interplanetary flight. We look at the deep-space interplanetary portion of the trajectory under rather idealized conditions. The two major performance indexes, the propellant mass and the flight time, are in conflict with one another for the following reason: any attempt at reducing the former causes an increase in the latter and vice versa. Therefore, it is natural to consider a compromise performance index involving the scaled values of the propellant mass and flight time weighted respectively by the compromise factor C and its complement 1-C. We use the compromise factor as the parameter of the one-parameter family of compromise trajectories. Analyses carried out with the sequential gradient-restoration algorithm for optimal control problems lead to results which can be highlighted as follows. Thrust profile. Generally speaking, the thrust profile of the compromise trajectory includes three subarcs: the first subarc is characterized by maximum thrust in conjunction with positive (upward) thrust direction; the second subarc is characterized by zero thrust (coasting flight); the third subarc is characterized by maximum thrust in conjunction with negative (downward) thrust direction. Effect of the compromise factor. As the compromise factor increases, the propellant mass decreases and the flight time increases; correspondingly, the following changes in the thrust profile take place: (a) the time lengths of the first and third subarcs (powered phases) decrease slightly, meaning that thrust application occurs for shorter duration; also, the average value of the thrust direction in the first and third subarcs decreases, implying higher efficiency of thrust application wrt the spacecraft energy level; as a result, the total propellant mass decreases; (b) the time length of the second subarc (coasting) increases considerably, resulting in total time increase. Minimum time trajectory. If C=0, the resulting minimum time trajectory has the following characteristics: (a) the time length of the coasting subarc reduces to zero and the three-subarc trajectory degenerates into a two-subarc trajectory; (b) maximum thrust is applied at all times and the thrust direction switches from upward to downward at midcourse. Minimum propellant mass trajectory. If C=1, the resulting minimum propellant mass trajectory has the following characteristics: (a) the thrust magnitude has a bang-zero-bang profile; (b) for the powered subarcs, the thrust direction is tangent to the flight path at all times.

Thrust Augmentation of a Turbojet Engine at Simulated Flight Conditions by Introduction of a Water-Alcohol Mixture into the Compressor

NASA Technical Reports Server (NTRS)

Useller, James W.; Auble, Carmon M.; Harvey, Ray W., Sr.

1952-01-01

An investigation was conducted at simulated high-altitude flight conditions to evaluate the use of compressor evaporative cooling as a means of turbojet-engine thrust augmentation. Comparison of the performance of the engine with water-alcohol injection at the compressor inlet, at the sixth stage of the compressor, and at the sixth and ninth stages was made. From consideration of the thrust increases achieved, the interstage injection of the coolant was considered more desirable preferred over the combined sixth- and ninth-stage injection because of its relative simplicity. A maximum augmented net-thrust ratio of 1.106 and a maximum augmented jet-thrust ratio of 1.062 were obtained at an augmented liquid ratio of 2.98 and an engine-inlet temperature of 80 F. At lower inlet temperatures (-40 to 40 F), the maximum augmented net-thrust ratios ranged from 1.040 to 1.076 and the maximum augmented jet-thrust ratios ranged from 1.027 to 1.048, depending upon the inlet temperature. The relatively small increase in performance at the lower inlet-air temperatures can be partially attributed to the inadequate evaporation of the water-alcohol mixture, but the more significant limitation was believed to be caused by the negative influence of the liquid coolant on engine- component performance. In general, it is concluded that the effectiveness of the injection of a coolant into the compressor as a means of thrust augmentation is considerably influenced by the design characteristics of the components of the engine being used.

Experimental Determination of Exhaust Gas Thrust, Special Report

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Voss, Fred

1940-01-01

This investigation presents the results of tests made on a radial engine to determine the thrust that can be obtained from the exhaust gas when discharged from separate stacks and when discharged from the collector ring with various discharge nozzles. The engine was provided with a propeller to absorb the power and was mounted on a test stand equipped with scales for measuring the thrust and engine torque. The results indicate that at full open throttle at sea level, for the engine tested, a gain in thrust horsepower of 18 percent using separate stacks, and 9.5 percent using a collector ring and discharge nozzle, can be expected at an air speed of 550 miles per hour.

Influence of Thrust Level on the Architecture and Optimal Working Process Parameters of a Small-scale Turbojet for UAV

NASA Astrophysics Data System (ADS)

Kuz`michev, V. S.; Filinov, E. P.; Ostapyuk, Ya A.

2018-01-01

This article describes how the thrust level influences the turbojet architecture (types of turbomachines that provide the maximum efficiency) and its working process parameters (turbine inlet temperature (TIT) and overall pressure ratio (OPR)). Functional gasdynamic and strength constraints were included, total mass of fuel and the engine required for mission and the specific fuel consumption (SFC) were considered optimization criteria. Radial and axial turbines and compressors were considered. The results show that as the engine thrust decreases, optimal values of working process parameters decrease too, and the regions of compromise shrink. Optimal engine architecture and values of working process parameters are suggested for turbojets with thrust varying from 100N to 100kN. The results show that for the thrust below 25kN the engine scale factor should be taken into the account, as the low flow rates begin to influence the efficiency of engine elements substantially.

Large-Scale Wind-Tunnel Tests of Exhaust Ingestion Due to Thrust Reversal on a Four-Engine Jet Transport during Ground Roll

NASA Technical Reports Server (NTRS)

Tolhurst, William H., Jr.; Hickey, David H.; Aoyagi, Kiyoshi

1961-01-01

Wind-tunnel tests have been conducted on a large-scale model of a swept-wing jet transport type airplane to study the factors affecting exhaust gas ingestion into the engine inlets when thrust reversal is used during ground roll. The model was equipped with four small jet engines mounted in nacelles beneath the wing. The tests included studies of both cascade and target type reversers. The data obtained included the free-stream velocity at the occurrence of exhaust gas ingestion in the outboard engine and the increment of drag due to thrust reversal for various modifications of thrust reverser configuration. Motion picture films of smoke flow studies were also obtained to supplement the data. The results show that the free-stream velocity at which ingestion occurred in the outboard engines could be reduced considerably, by simple modifications to the reversers, without reducing the effective drag due to reversed thrust.

Thermal Barrier and Protective Coatings to Improve the Durability of a Combustor Under a Pulse Detonation Engine Environment

NASA Technical Reports Server (NTRS)

Ghosn, Louis J.; Zhu, Dongming

2008-01-01

Pulse detonation engine (PDE) concepts are receiving increasing attention for future aeronautic propulsion applications, due to their potential thermodynamic cycle efficiency and higher thrust to density ratio that lead to the decrease in fuel consumption. But the resulting high gas temperature and pressure fluctuation distributions at high frequency generated with every detonation are viewed to be detrimental to the combustor liner material. Experimental studies on a typical metal combustion material exposed to a laser simulated pulse heating showed extensive surface cracking. Coating of the combustor materials with low thermal conductivity ceramics is shown to protect the metal substrate, reduce the thermal stresses, and hence increase the durability of the PDE combustor liner material. Furthermore, the temperature fluctuation and depth of penetration is observed to decrease with increasing the detonation frequency. A crack propagation rate in the coating is deduced by monitoring the variation of the coating apparent thermal conductivity with time that can be utilized as a health monitoring technique for the coating system under a rapid fluctuating heat flux.

Exhaust-stack nozzle area and shape for individual cylinder exhaust-gas jet-propulsion system

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Turner, Richard; Voss, Fred; Humble, Leroy V

1943-01-01

This report presents the results of an investigation conducted on the effect of exhaust-stack nozzle area, shape, and length on engine power, jet thrust, and gain in net thrust (engine propeller plus jet). Single-cylinder engine data were obtained using three straight stacks 25, 44, and 108 inches in length; an S-shaped stack, a 90 degree bend, a 180 degree bend, and a short straight stack having a closed branch faired into it. Each stack was fitted with nozzles varying in exit area from 0.91 square inch to the unrestricted area of the stack of 4.20 square inches. The engine was generally operated over a range of engine speeds from 1300 to 2100 r.p.m, inlet-manifold pressures from 22 to 30 inches of mercury absolute, and a fuel-air ratio of 0.08. The loss in engine power, the jet thrust, and the gain in net thrust are correlated in terms of several simple parameters. An example is given for determining the optimum nozzle area and the overall net thrust.

NASA's Nuclear Thermal Propulsion Project

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Mitchell, Doyce P.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Gerrish, Harold P.; Doughty, Glen; Belvin, Anthony; Clement, Steven; Borowski, Stanley K.;

The use of laterally vectored thrust to counter thrust asymmetry in a tactical jet aircraft

NASA Technical Reports Server (NTRS)

1983-01-01

A nonlinear, six degree-of-freedom flight simulator for a twin engine tactical jet was built on a hybrid computer to investigate lateral vectoring of the remaining thrust component for the case of a single engine failure at low dynamic pressures. Aircraft control was provided by an automatic controller rather than a pilot, and thrust vector control was provided by an open-loop controller that deflected a vane (located on the periphery of each exhaust jet and normally streamlined for noninterference with the flow). Lateral thrust vectoring decreased peak values of lateral control deflections, eliminated the requirement for steady-state lateral aerodynamic control deflections, and decreased the amount of altitude lost for a single engine failure.

Engineering Research and Development and Technology thrust area report FY92

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

Langland, R.T.; Minichino, C.

1993-03-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the technical staff and the technology needed to support current and future LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Program has two important goals: (1) to identify key technologies and (2) to conduct high-quality work to enhance our capabilities in these key technologies. To help focus our efforts, we identify technology thrust areas and select technical leaders for each area. The thrust areas are integrated engineering activities and, rather than being based on individual disciplines, theymore » are staffed by personnel from Electronics Engineering, Mechanical Engineering, and other LLNL organizations, as appropriate. The thrust area leaders are expected to establish strong links to LLNL program leaders and to industry; to use outside and inside experts to review the quality and direction of the work; to use university contacts to supplement and complement their efforts; and to be certain that we are not duplicating the work of others. This annual report, organized by thrust area, describes activities conducted within the Program for the fiscal year 1992. Its intent is to provide timely summaries of objectives, theories, methods, and results. The nine thrust areas for this fiscal year are: Computational Electronics and Electromagnetics; Computational Mechanics; Diagnostics and Microelectronics; Emerging Technologies; Fabrication Technology; Materials Science and Engineering; Microwave and Pulsed Power; Nondestructive Evaluation; and Remote Sensing and Imaging, and Signal Engineering.« less

Quiet Clean Short-Haul Experimental Engine (QCSEE) acoustic and aerodynamic tests on a scale model over-the-wing thrust reverser and forward thrust nozzle

NASA Technical Reports Server (NTRS)

Stimpert, D. L.

1978-01-01

An acoustic and aerodynamic test program was conducted on a 1/6.25 scale model of the Quiet, Clean, Short-Haul Experimental Engine (QCSEE) forward thrust over-the-wing (OTW) nozzle and OTW thrust reverser. In reverse thrust, the effect of reverser geometry was studied by parametric variations in blocker spacing, blocker height, lip angle, and lip length. Forward thrust nozzle tests determined the jet noise levels of the cruise and takeoff nozzles, the effect of opening side doors to achieve takeoff thrust, and scrubbing noise of the cruise and takeoff jet on a simulated wing surface. Velocity profiles are presented for both forward and reverse thrust nozzles. An estimate of the reverse thrust was made utilizing the measured centerline turning angle.

The Strutjet Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

NASA Orbit Transfer Rocket Engine Technology Program

NASA Technical Reports Server (NTRS)

1984-01-01

The advanced expander cycle engine with a 15,000 lb thrust level and a 6:1 mixture ratio and optimized performance was used as the baseline for a design study of the hydrogen/oxgyen propulsion system for the orbit transfer vehicle. The critical components of this engine are the thrust chamber, the turbomachinery, the extendible nozzle system, and the engine throttling system. Turbomachinery technology is examined for gears, bearing, seals, and rapid solidification rate turbopump shafts. Continuous throttling concepts are discussed. Components of the OTV engine described include the thrust chamber/nozzle assembly design, nozzles, the hydrogen regenerator, the gaseous oxygen heat exchanger, turbopumps, and the engine control valves.

Development of Modeling Approaches for Nuclear Thermal Propulsion Test Facilities

NASA Technical Reports Server (NTRS)

Jones, Daniel R.; Allgood, Daniel C.; Nguyen, Ke

2014-01-01

High efficiency of rocket propul-sion systems is essential for humanity to venture be-yond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rock-ets with relatively high thrust and twice the efficiency of the Space Shuttle Main Engine. NASA is in the pro-cess of developing a new NTP engine, and is evaluat-ing ground test facility concepts that allow for the thor-ough testing of NTP devices. NTP engine exhaust, hot gaseous hydrogen, is nominally expected to be free of radioactive byproducts from the nuclear reactor; how-ever, it has the potential to be contaminated due to off-nominal engine reactor performance. Several options are being investigated to mitigate this hazard potential with one option in particular that completely contains the engine exhaust during engine test operations. The exhaust products are subsequently disposed of between engine tests. For this concept (see Figure 1), oxygen is injected into the high-temperature hydrogen exhaust that reacts to produce steam, excess oxygen and any trace amounts of radioactive noble gases released by off-nominal NTP engine reactor performance. Water is injected to condense the potentially contaminated steam into water. This water and the gaseous oxygen (GO2) are subsequently passed to a containment area where the water and GO2 are separated into separate containment tanks.

Characterization of Low-Frequency Combustion Stability of the Fastrac Engine

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Jones, Preston (Technical Monitor)

2002-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

The feasibility of potential reusable thrust chamber concepts is studied. Propellant condidates were examined and analytically combined with potential cooling schemes. A data base of engine data which would assist in a configuration selection was produced. The data base verification was performed by the demonstration of a thrust chamber of a selected coolant scheme design. A full scale insulated columbium thrust chamber was used for propellant coolant configurations. Combustion stability of the injectors and a reduced size thrust chamber were experimentally verified as proof of concept demonstrations of the design and study results.

Operationalizing Special Operations Aviation in Indonesia

DTIC Science & Technology

2006-12-15

special operations forces Builder: Lockheed Power Plant: Four Allison T56 -A-15 turboprop engines Thrust: 4,910 shaft horsepower each engine...Builder: Lockheed Power Plant: Four Allison T56 -A-15 turboprop engines Thrust: 4,910 shaft horsepower each engine Length: 98 feet, 9 inches (30.09

Isomer Energy Source for Space Propulsion Systems

DTIC Science & Technology

2004-03-01

1,590 Engine F/W (no shield) 3.4 5.0 20.0 A similar core design replacing the fission fuel with the isomer 178Hfm2 is the starting point for this...particles interact and collide with other atoms in the fuel material, reactor core , or coolant, their energy can be transferred to thermal energy...thrust (44). The program produced several reactors that made it all the way through the testing stages of development . The reactors used uranium-235

Pulse Detonation Engine Thrust Tube Heat Exchanger for Flash Vaporization and Supercritical Heating of JP-8

DTIC Science & Technology

2005-03-01

47 Figure 21. Construction of the long heat exchanger with helical rod welded in place.... 48 Figure 22. Heat exchanger...not at a temperature at or above the dew point temperature of the mixture, some of the fuel in the mixture will re- condense . The concept of...diao (25) Where kamb = Thermal conductivity of the air [W/(m-K)] Nufc = Nusselt number for free convection The Nussult number

Low-thrust chemical propulsion system propellant expulsion and thermal conditioning study

NASA Technical Reports Server (NTRS)

Merino, F.; Wakabayashi, I.; Pleasant, R. L.; Hill, M.

1982-01-01

Thermal conditioning systems for satisfying engine net positive suction pressure (NPSP) requirements, and propellant expulsion systems for achieving propellant dump during a return-to-launch site (RTLS) abort were studied for LH2/LO2 and LCH4/LO2 upper stage propellant combinations. A state-of-the-art thermal conditioning system employing helium injection beneath the liquid surface shows the lowest weight penalty for LO2 and LCH4. A technology system incorporating a thermal subcooler (heat exchanger) for engine NPSP results in the lowest weight penalty for the LH2 tank. A preliminary design of two state-of-the-art and two new technology systems indicates a weight penalty difference too small to warrant development of a LH2 thermal subcooler. Analysis results showed that the LH2/LO2 propellant expulsion system is optimized for maximum dump line diameters, whereas the LCH4/LO2 system is optimized for minimum dump line diameter (LCH4) and maximum dump line diameter (LO2). The primary uncertainty is the accurate determination of two-phase flow rates through the dump system; experimentation is not recommended because this uncertainty is not considered significant.

A Combustion Research Facility for Testing Advanced Materials for Space Applications

NASA Technical Reports Server (NTRS)

Bur, Michael J.

2003-01-01

The test facility presented herein uses a groundbased rocket combustor to test the durability of new ceramic composite and metallic materials in a rocket engine thermal environment. A gaseous H2/02 rocket combustor (essentially a ground-based rocket engine) is used to generate a high temperature/high heat flux environment to which advanced ceramic and/or metallic materials are exposed. These materials can either be an integral part of the combustor (nozzle, thrust chamber etc) or can be mounted downstream of the combustor in the combustor exhaust plume. The test materials can be uncooled, water cooled or cooled with gaseous hydrogen.

14 CFR 33.8 - Selection of engine power and thrust ratings.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Selection of engine power and thrust ratings. 33.8 Section 33.8 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES General § 33.8 Selection of engine power and...

14 CFR 33.8 - Selection of engine power and thrust ratings.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Selection of engine power and thrust ratings. 33.8 Section 33.8 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES General § 33.8 Selection of engine power and...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

Nuclear Thermal Propulsion Mars Mission Systems Analysis and Requirements Definition

NASA Technical Reports Server (NTRS)

Mulqueen, Jack; Chiroux, Robert C.; Thomas, Dan; Crane, Tracie

2007-01-01

This paper describes the Mars transportation vehicle design concepts developed by the Marshall Space Flight Center (MSFC) Advanced Concepts Office. These vehicle design concepts provide an indication of the most demanding and least demanding potential requirements for nuclear thermal propulsion systems for human Mars exploration missions from years 2025 to 2035. Vehicle concept options vary from large "all-up" vehicle configurations that would transport all of the elements for a Mars mission on one vehicle. to "split" mission vehicle configurations that would consist of separate smaller vehicles that would transport cargo elements and human crew elements to Mars separately. Parametric trades and sensitivity studies show NTP stage and engine design options that provide the best balanced set of metrics based on safety, reliability, performance, cost and mission objectives. Trade studies include the sensitivity of vehicle performance to nuclear engine characteristics such as thrust, specific impulse and nuclear reactor type. Tbe associated system requirements are aligned with the NASA Exploration Systems Mission Directorate (ESMD) Reference Mars mission as described in the Explorations Systems Architecture Study (ESAS) report. The focused trade studies include a detailed analysis of nuclear engine radiation shield requirements for human missions and analysis of nuclear thermal engine design options for the ESAS reference mission.

14 CFR Appendix I to Part 25 - Installation of an Automatic Takeoff Thrust Control System (ATTCS)

Code of Federal Regulations, 2010 CFR

2010-01-01

...) This appendix specifies additional requirements for installation of an engine power control system that... crew to increase thrust or power. I25.2Definitions. (a) Automatic Takeoff Thrust Control System (ATTCS... mechanical and electrical, that sense engine failure, transmit signals, actuate fuel controls or power levers...

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.

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.

Nuclear Thermal Propulsion (NTP): A Proven Growth Technology for Human NEO/Mars Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The nuclear thermal rocket (NTR) represents the next "evolutionary step" in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine s reactor core. Using an "expander" cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (Isp) values of 900 seconds (s) or more - twice that of today s best chemical rockets. From 1955 - 1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability - all the requirements needed for a human Mars mission. Ceramic metal "cermet" fuel was pursued as well, as a backup option. The NTR also has significant "evolution and growth" capability. Configured as a "bimodal" system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen "afterburner" nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program - the 25,000 lbf (25 klbf) "Pewee" engine is sufficient when used in a clustered engine arrangement. The "Copernicus" crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA s current activities and future plans for NTP development that include system-level Technology Demonstrations - specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

Dynamic Characterization of an Inflatable Concentrator for Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

Leigh, Larry; Hamidzadeh, Hamid; Tinker, Michael L.; Rodriguez, Pedro I. (Technical Monitor)

2001-01-01

An inflatable structural system that is a technology demonstrator for solar thermal propulsion and other applications is characterized for structural dynamic behavior both experimentally and computationally. The inflatable structure is a pressurized assembly developed for use in orbit to support a Fresnel lens or inflatable lenticular element for focusing sunlight into a solar thermal rocket engine. When the engine temperature reaches a pre-set level, the propellant is injected into the engine, absorbs heat from an exchanger, and is expanded through the nozzle to produce thrust. The inflatable structure is a passively adaptive system in that a regulator and relief valve are utilized to maintain pressure within design limits during the full range of orbital conditions. Modeling and test activities are complicated by the fact that the polyimide film material used for construction of the inflatable is nonlinear, with modulus varying as a function of frequency, temperature, and level of excitation. Modal vibration testing and finite element modeling are described in detail in this paper. The test database is used for validation and modification of the model. This work is highly significant because of the current interest in inflatable structures for space application, and because of the difficulty in accurately modeling such systems.

Titanium honeycomb acoustic lining structural and thermal test report. [for acoustic tailpipe for JT8D engine

NASA Technical Reports Server (NTRS)

Joynes, D.; Balut, J. P.

1974-01-01

The results are presented of static, fatigue and thermal testing of titanium honeycomb acoustic panels representing the acoustic tailpipe for the Pratt and Whitney Aircraft JT8D Refan engine which is being studied for use on the Boeing 727-200 airplane. Test specimens represented the engine and tailpipe flange joints, the rail to which the thrust reverser is attached and shear specimens of the tailpipe honeycomb. Specimens were made in four different batches with variations in configuration, materials and processes in each. Static strength of all test specimens exceeded the design ultimate load requirements. Fatigue test results confirmed that aluminum brazed titanium, as used in the Refan tailpipe design, meets the fatigue durability objectives. Quality of welding was found to be critical to life, with substandard welding failing prematurely, whereas welding within the process specification exceeded the panel skin life. Initial fatigue testing used short grip length bolts which failed prematurely. These were replaced with longer bolts and subsequent testing demonstrated the required life. Thermal tests indicate that perforated skin acoustic honeycomb has approximately twice the heat transfer of solid skin honeycomb.

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

DTIC Science & Technology

2009-01-01

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

Comparison of Performance and Component Frontal Areas of Hypothetical Two-spool and One-spool Turbojet Engines

NASA Technical Reports Server (NTRS)

Dugan, James F , Jr

1956-01-01

For constant-mechanical-speed operation, the two-spool thrust values are as great as or greater than the one-spool thrust values over the entire flight range considered, while the specific fuel consumption for the two engines agrees within 1 percent. The maximum difference in thrust occurs at Mach 2.8 in the stratosphere, where the two-spool thrust advantage is about 9 percent for operation with the after burning.

Static Performance of a Wing-Mounted Thrust Reverser Concept

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

1998-01-01

An experimental investigation was conducted in the Jet-Exit Test Facility at NASA Langley Research Center to study the static aerodynamic performance of a wing-mounted thrust reverser concept applicable to subsonic transport aircraft. This innovative engine powered thrust reverser system is designed to utilize wing-mounted flow deflectors to produce aircraft deceleration forces. Testing was conducted using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0, a supercritical left-hand wing section attached via a pylon, and wing-mounted flow deflectors attached to the wing section. Geometric variations of key design parameters investigated for the wing-mounted thrust reverser concept included flow deflector angle and chord length, deflector edge fences, and the yaw mount angle of the deflector system (normal to the engine centerline or parallel to the wing trailing edge). All tests were conducted with no external flow and high pressure air was used to simulate core and fan engine exhaust flows. Test results indicate that the wing-mounted thrust reverser concept can achieve overall thrust reverser effectiveness levels competitive with (parallel mount), or better than (normal mount) a conventional cascade thrust reverser system. By removing the thrust reverser system from the nacelle, the wing-mounted concept offers the nacelle designer more options for improving nacelle aero dynamics and propulsion-airframe integration, simplifying nacelle structural designs, reducing nacelle weight, and improving engine maintenance access.

Method and apparatus for rapid thrust increases in a turbofan engine

NASA Technical Reports Server (NTRS)

Cornett, J. E.; Corley, R. C.; Fraley, T. O.; Saunders, A. A., Jr. (Inventor)

1980-01-01

Upon a landing approach, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to thereby increase the fuel flow and compressor speed in order to maintain fan speed and thrust. This running of the compressor at an off-design speed substantially reduces the time required to subsequently advance the engine speed to the takeoff thrust level by advancing the throttle and opening the compressor stators.

Converting the ISS to an Earth-Moon Transport System Using Nuclear Thermal Propulsion

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

Paniagua, John; Maise, George; Powell, James

2008-01-21

Using Nuclear Thermal Propulsion (NTP), the International Space Station (ISS) can be placed into a cyclic orbit between the Earth and the Moon for 2-way transport of personnel and supplies to a permanent Moon Base. The ISS cycler orbit apogees 470,000 km from Earth, with a period of 13.66 days. Once a month, the ISS would pass close to the Moon, enabling 2-way transport between it and the surface using a lunar shuttle craft. The lunar shuttle craft would land at a desired location on the surface during a flyby and return to the ISS during a later flyby. Atmore » Earth perigee 7 days later at 500 km altitude, there would be 2-way transport between it and Earth's surface using an Earth shuttle craft. The docking Earth shuttle would remain attached to the ISS as it traveled towards the Moon, while personnel and supplies transferred to a lunar shuttle spacecraft that would detach and land at the lunar base when the ISS swung around the Moon. The reverse process would be carried out to return personnel and materials from the Moon to the Earth. The orbital mechanics for the ISS cycle are described in detail. Based on the full-up mass of 400 metric tons for the ISS, an ISP of 900 seconds, and a delta V burn of 3.3 km/sec to establish the orbit, 200 metric tons of liquid H-2 propellant would be required. The 200 metric tons could be stored in 3 tanks, each 8 meters in diameter and 20 meters in length. An assembly of 3 MITEE NTP engines would be used, providing redundancy if an engine were to fail. Two different MITEE design options are described. Option 1 is an 18,000 Newton, 100 MW engine with a thrust to weight ratio of 6.6/1; Option 2 is a 180,000 Newton, 1000 MW engine with a thrust to weight ratio of 23/1. Burn times to establish the orbit are {approx}1 hour for the large 3 engine assembly, and 10 hours for the small 3 engine assembly. Both engines would use W-UO2 cermet fuel at {approx}2750 K which has demonstrated the capability to operate for at least 50 hours in 2750 K hydrogen with only a minor loss of fuel material. The small engine is favored because of its lower weight. The total system weight of the small 3 engine assembly is {approx}12 metric tons, including engine, controls, pumps, and neutron and gamma shields. After their main thrust operation, the NTP engines would shut down, with periodic successive smaller delta V burns as required to fine-tune the cycler orbit. Radiation dosages to personnel, both during operation and after shutdown, are much smaller than those from the cosmic ray background.« less

Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

NASA Technical Reports Server (NTRS)

Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

1977-01-01

This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 transport aircraft approaches its first landing under engine power only on Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 approaches the first landing ever of a transport aircraft under engine power only on Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when it normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

Predicted performance of an integrated modular engine system

NASA Technical Reports Server (NTRS)

Binder, Michael; Felder, James L.

1993-01-01

Space vehicle propulsion systems are traditionally comprised of a cluster of discrete engines, each with its own set of turbopumps, valves, and a thrust chamber. The Integrated Modular Engine (IME) concept proposes a vehicle propulsion system comprised of multiple turbopumps, valves, and thrust chambers which are all interconnected. The IME concept has potential advantages in fault-tolerance, weight, and operational efficiency compared with the traditional clustered engine configuration. The purpose of this study is to examine the steady-state performance of an IME system with various components removed to simulate fault conditions. An IME configuration for a hydrogen/oxygen expander cycle propulsion system with four sets of turbopumps and eight thrust chambers has been modeled using the Rocket Engine Transient Simulator (ROCETS) program. The nominal steady-state performance is simulated, as well as turbopump thrust chamber and duct failures. The impact of component failures on system performance is discussed in the context of the system's fault tolerant capabilities.

Background and principles of throttles-only flight control

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.

1995-01-01

There have been many cases in which the crew of a multi-engine airplane had to use engine thrust for emergency flight control. Such a procedure is very difficult, because the propulsive control forces are small, the engine response is slow, and airplane dynamics such as the phugoid and dutch roll are difficult to damp with thrust. In general, thrust increases are used to climb, thrust decreases to descend, and differential thrust is used to turn. Average speed is not significantly affected by changes in throttle setting. Pitch control is achieved because of pitching moments due to speed changes, from thrust offset, and from the vertical component of thrust. Roll control is achieved by using differential thrust to develop yaw, which, through the normal dihedral effect, causes a roll. Control power in pitch and roll tends to increase as speed decreases. Although speed is not controlled by the throttles, configuration changes are often available (lowering gear, flaps, moving center-of-gravity) to change the speed. The airplane basic stability is also a significant factor. Fuel slosh and gyroscopic moments are small influences on throttles-only control. The background and principles of throttles-only flight control are described.

A study to estimate and compare the total particulate matter emission indices (EIN) between traditional jet fuel and two blends of Jet A/Camelina biofuel used in a high by-pass turbofan engine: A case study of Honeywell TFE-109 engine

NASA Astrophysics Data System (ADS)

Shila, Jacob Joshua Howard

The aviation industry is expected to grow at an annual rate of 5% until the year 2031 according to Boeing Outlook Report of 2012. Although the aerospace manufacturers have introduced new aircraft and engines technologies to reduce the emissions generated by aircraft engines, about 15% of all aircraft in 2032 will be using the older technologies. Therefore, agencies such as the National Aeronautics and Astronautics Administration (NASA), Federal Aviation Administration (FAA), the Environmental Protection Agency (EPA) among others together with some academic institutions have been working to characterize both physical and chemical characteristics of the aircraft particulate matter emissions to further understand their effects to the environment. The International Civil Aviation Organization (ICAO) is also working to establish an inventory with Particulate Matter emissions for all the aircraft turbine engines for certification purposes. This steps comes as a result of smoke measurements not being sufficient to provide detailed information on the effects of Particulate Matter (PM) emissions as far as the health and environmental concerns. The use of alternative fuels is essential to reduce the impacts of emissions released by Jet engines since alternative aviation fuels have been studied to lower particulate matter emissions in some types of engines families. The purpose of this study was to determine whether the emission indices of the biofuel blended fuels were lower than the emission indices of the traditional jet fuel at selected engine thrust settings. The biofuel blends observed were 75% Jet A-25% Camelina blend biofuel, and 50% Jet A-50% Jet A blend biofuel. The traditional jet fuel in this study was the Jet A fuel. The results of this study may be useful in establishing a baseline for aircraft engines' PM inventory. Currently the International Civil Aviation Organization (ICAO) engines emissions database contains only gaseous emissions data for only the TFE 731 and JT15D engines' families as representatives of other engines with rated thrust of 6000 pounds or below. The results of this study may be used to add to the knowledge of PM emission data that has been collected in other research studies. This study was quantitative in nature. Three factors were designated which were the types of fuels studied. The TFE-109 turbofan engine was the experimental subject. The independent variable was the engine thrust setting while the response variable was the emission index. Four engine runs were conducted for each fuel. In each engine run, four engine thrust settings were observed. The four engine thrust levels were 10%, 30%, 85%, and 100% rated thrusts levels. Therefore, for each engine thrust settings, there four replicates. The experiments were conducted using a TFE-109 engine test cell located in the Niswonger Aviation Technology building at the Purdue University Airport. The testing facility has the capability to conduct the aircraft PM emissions tests. Due to the equipment limitations, the study was limited to observe total PM emissions instead of specifically measuring the non-volatile PM emissions. The results indicate that the emissions indices of the blended biofuels were not statistically significantly lower compared to the emissions of the traditional jet fuel at rated thrust levels of 100% and 85% of TFE-109 turbofan engine. However, the emission indices for the 50%Jet A - 50%Camelina biofuel blend were statistically significantly lower compared to the emission indices of the 100% Jet A fuel at 10% and 30% engine rated thrusts levels of TFE-109 engine. The emission indices of the 50%-50% biofuel blend were lower by reductions of 15% and 17% at engine rated thrusts of 10% and 30% respectively compared to the emissions indices of the traditional jet fuel at the same engine thrust levels. Experimental modifications in future studies may provide estimates of the emissions indices range for this particular engine these estimates may be used to estimate the levels of PM emissions for other similar engines. Additional measurements steps such as heating of the sampling line, sampling dilution application, sampling line loss estimates, and calculations of the sampling line PM residence times will also be useful future results.

Mini-MITEE: Ultra Small, Ultra Light NTP Engines for Robotic Science and Manned Exploration Missions

NASA Astrophysics Data System (ADS)

Powell, James; Maise, George; Paniagua, John

2006-01-01

A compact, ultra lightweight Nuclear Thermal Propulsion (NTP) engine design is described with the capability to carry out a wide range of unique and important robotic science missions that are not possible using chemical or Nuclear Electric Propulsion (NEP). The MITEE (MInature ReacTor EnginE) reactor uses hydrogeneous moderator, such as solid lithium-7 hydride, and high temperature cermet tungsten/UO2 nuclear fuel. The reactor is configured as a modular pressure tube assembly, with each pressure tube containing an outer annual shell of moderator with an inner annular region of W/UO2 cermet fuel sheets. H2 propellant flows radially inwards through the moderator and fuel regions, exiting at ~3000 K into a central channel that leads to a nozzle at the end of the pressure tube. Power density in the fuel region is 10 to 20 megawatts per liter, depending on design, producing a thrust output on the order of 15,000 Newtons and an Isp of ~1000 seconds. 3D Monte Carlo neutronic analyses are described for MITEE reactors utilizing various fissile fuel options (U-235, U-233, and Am242m) and moderators (7LiH and BeH2). Reactor mass ranges from a maximum of 100 kg for the 7LiH/U-235 option to a minimum of 28 kg for the BeH2/Am-242 m option. Pure thrust only and bi-modal (thrust plus electric power generation) MITEE designs are described. Potential unique robotic science missions enabled by the MITEE engine are described, including landing on Europa and exploring the ice sheet interior with return of samples to Earth, hopping to and exploring multiple sites on Mars, unlimited ramjet flight in the atmospheres of Jupiter, Saturn, Uranus, and Neptune and landing on, and sample return from Pluto.

Mini-MITEE: Ultra Small, Ultra Light NTP Engines for Robotic Science and Manned Exploration Missions

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

Powell, James; Maise, George; Paniagua, John

2006-01-20

A compact, ultra lightweight Nuclear Thermal Propulsion (NTP) engine design is described with the capability to carry out a wide range of unique and important robotic science missions that are not possible using chemical or Nuclear Electric Propulsion (NEP). The MITEE (MInature ReacTor EnginE) reactor uses hydrogeneous moderator, such as solid lithium-7 hydride, and high temperature cermet tungsten/UO2 nuclear fuel. The reactor is configured as a modular pressure tube assembly, with each pressure tube containing an outer annual shell of moderator with an inner annular region of W/UO2 cermet fuel sheets. H2 propellant flows radially inwards through the moderator andmore » fuel regions, exiting at {approx}3000 K into a central channel that leads to a nozzle at the end of the pressure tube. Power density in the fuel region is 10 to 20 megawatts per liter, depending on design, producing a thrust output on the order of 15,000 Newtons and an Isp of {approx}1000 seconds. 3D Monte Carlo neutronic analyses are described for MITEE reactors utilizing various fissile fuel options (U-235, U-233, and Am242m) and moderators (7LiH and BeH2). Reactor mass ranges from a maximum of 100 kg for the 7LiH/U-235 option to a minimum of 28 kg for the BeH2/Am-242 m option. Pure thrust only and bi-modal (thrust plus electric power generation) MITEE designs are described. Potential unique robotic science missions enabled by the MITEE engine are described, including landing on Europa and exploring the ice sheet interior with return of samples to Earth, hopping to and exploring multiple sites on Mars, unlimited ramjet flight in the atmospheres of Jupiter, Saturn, Uranus, and Neptune and landing on, and sample return from Pluto.« less

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

NASA Technical Reports Server (NTRS)

Betts, Erin M.; Frederick, Robert A., Jr.

2010-01-01

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Maximum thrust mode evaluation

NASA Technical Reports Server (NTRS)

Orme, John S.; Nobbs, Steven G.

1995-01-01

Measured reductions in acceleration times which resulted from the application of the F-15 performance seeking control (PSC) maximum thrust mode during the dual-engine test phase is presented as a function of power setting and flight condition. Data were collected at altitudes of 30,000 and 45,000 feet at military and maximum afterburning power settings. The time savings for the supersonic acceleration is less than at subsonic Mach numbers because of the increased modeling and control complexity. In addition, the propulsion system was designed to be optimized at the mid supersonic Mach number range. Recall that even though the engine is at maximum afterburner, PSC does not trim the afterburner for the maximum thrust mode. Subsonically at military power, time to accelerate from Mach 0.6 to 0.95 was cut by between 6 and 8 percent with a single engine application of PSC, and over 14 percent when both engines were optimized. At maximum afterburner, the level of thrust increases were similar in magnitude to the military power results, but because of higher thrust levels at maximum afterburner and higher aircraft drag at supersonic Mach numbers the percentage thrust increase and time to accelerate was less than for the supersonic accelerations. Savings in time to accelerate supersonically at maximum afterburner ranged from 4 to 7 percent. In general, the maximum thrust mode has performed well, demonstrating significant thrust increases at military and maximum afterburner power. Increases of up to 15 percent at typical combat-type flight conditions were identified. Thrust increases of this magnitude could be useful in a combat situation.

Evaluation of a ducted-fan power plant designed for high output and good cruise fuel economy

NASA Technical Reports Server (NTRS)

Behun, M; Rom, F E; Hensley, R V

1950-01-01

Theoretical analysis of performance of a ducted-fan power plant designed both for high-output, high-altitude operation at low supersonic Mach numbers and for good fuel economy at lower fight speeds is presented. Performance of ducted fan is compared with performance (with and without tail-pipe burner) of two hypothetical turbojet engines. At maximum power, the ducted fan has propulsive thrust per unit of frontal area between thrusts obtained by turbojet engines with and without tail-pipe burners. At cruise, the ducted fan obtains lowest thrust specific fuel consumption. For equal maximum thrusts, the ducted fan obtains cruising flight duration and range appreciably greater than turbojet engines.

Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

Quiet Clean Short-haul Experimental Engine (QCSEE) under-the-wing engine simulation report

NASA Technical Reports Server (NTRS)

1977-01-01

Hybrid computer simulations of the under-the-wing engine were constructed to develop the dynamic design of the controls. The engine and control system includes a variable pitch fan and a digital electronic control. Simulation results for throttle bursts from 62 to 100 percent net thrust predict that the engine will accelerate 62 to 95 percent net thrust in one second.

Propulsion Systems Panel

NASA Technical Reports Server (NTRS)

Bianca, Carmelo J.; Miner, Robert; Johnston, Lawrence M.; Bruce, R.; Dennies, Daniel P.; Dickenson, W.; Dreshfield, Robert; Karakulko, Walt; Mcgaw, Mike; Munafo, Paul M.

1993-01-01

Topics addressed are: (1) cryogenic tankage; (2) launch vehicle TPS/insulation; (3) durable passive thermal control devices and/or coatings; (4) development and characterization of processing methods to reduce anisotropy of material properties in Al-Li; (5) durable thermal protection system (TPS); (6) unpressurized Al-Li structures (interstages, thrust structures); (7) near net shape sections; (8) pressurized structures; (9) welding and joining; (10) micrometeoroid and debris hypervelocity shields; (11) state-of-the-art shell buckling structure optimizer program to serve as a rapid design tool; (12) test philosophy; (13) reduced load cycle time; (14) structural analysis methods; (15) optimization of structural criteria; and (16) develop an engineering approach to properly trade material and structural concepts selection, fabrication, facilities, and cost.

Propulsion Systems Panel

NASA Astrophysics Data System (ADS)

Bianca, Carmelo J.; Miner, Robert; Johnston, Lawrence M.; Bruce, R.; Dennies, Daniel P.; Dickenson, W.; Dreshfield, Robert; Karakulko, Walt; McGaw, Mike; Munafo, Paul M.

1993-02-01

Topics addressed are: (1) cryogenic tankage; (2) launch vehicle TPS/insulation; (3) durable passive thermal control devices and/or coatings; (4) development and characterization of processing methods to reduce anisotropy of material properties in Al-Li; (5) durable thermal protection system (TPS); (6) unpressurized Al-Li structures (interstages, thrust structures); (7) near net shape sections; (8) pressurized structures; (9) welding and joining; (10) micrometeoroid and debris hypervelocity shields; (11) state-of-the-art shell buckling structure optimizer program to serve as a rapid design tool; (12) test philosophy; (13) reduced load cycle time; (14) structural analysis methods; (15) optimization of structural criteria; and (16) develop an engineering approach to properly trade material and structural concepts selection, fabrication, facilities, and cost.

Safe, Affordable, Nuclear Thermal Propulsion Systems

NASA Technical Reports Server (NTRS)

Houts, M. G.; Kim, T.; Emrich, W. J.; Hickman, R. R.; Broadway, J. W.; Gerrish, H. P.; Doughty, G. E.

2014-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced Nuclear Electric Propulsion (NEP).

Nuclear Thermal Propulsion for Advanced Space Exploration

NASA Technical Reports Server (NTRS)

Houts, M. G.; Borowski, S. K.; George, J. A.; Kim, T.; Emrich, W. J.; Hickman, R. R.; Broadway, J. W.; Gerrish, H. P.; Adams, R. B.

2012-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced Nuclear Electric Propulsion (NEP).

Mariner Venus/Mercury 1973 rocket engine assembly

NASA Technical Reports Server (NTRS)

Snoke, D. R.; Williams, R. S.

1972-01-01

The fabrication and test of rocket engine assemblies (REA) for Mariner Venus/Mercury 1973 are reported. The fabrication, assembly and flight acceptance test of seven REA's including the type approval test of one engine and fabrication of one additional kit consisting of detail parts for an engine ready for catalyst loading are presented. The MV/M '73 REA which is a nominal 51 lbs thrust monopropellant engine is described. Under steady state operation the specific impulse is not less than 228 lb-sec at 55 lb and 218.5 lb-sec at 10 lb thrust varying linearly between these limits. The characteristic velocity is not less than 4100 ft/sec at any thrust level.

On spacecraft maneuvers control subject to propellant engine modes.

PubMed

Mazinan, A H

2015-09-01

The paper attempts to address a new control approach to spacecraft maneuvers based upon the modes of propellant engine. A realization of control strategy is now presented in engine on mode (high thrusts as well as further low thrusts), which is related to small angle maneuvers and engine off mode (specified low thrusts), which is also related to large angle maneuvers. There is currently a coarse-fine tuning in engine on mode. It is shown that the process of handling the angular velocities are finalized via rate feedback system in engine modes, where the angular rotations are controlled through quaternion based control (QBCL)strategy in engine off mode and these ones are also controlled through an optimum PID (OPIDH) strategy in engine on mode. Copyright © 2015 ISA. Published by Elsevier Ltd. All rights reserved.

Solar Thermal Upper Stage Cryogen System Engineering Checkout Test

NASA Technical Reports Server (NTRS)

Olsen, A. D; Cady, E. C.; Jenkins, D. S.

1999-01-01

The Solar Thermal Upper Stage technology (STUSTD) program is a solar thermal propulsion technology program cooperatively sponsored by a Boeing led team and by NASA MSFC. A key element of its technology program is development of a liquid hydrogen (LH2) storage and supply system which employs multi-layer insulation, liquid acquisition devices, active and passive thermodynamic vent systems, and variable 40W tank heaters to reliably provide near constant pressure H2 to a solar thermal engine in the low-gravity of space operation. The LH2 storage and supply system is designed to operate as a passive, pressure fed supply system at a constant pressure of about 45 psia. During operation of the solar thermal engine over a small portion of the orbit the LH2 storage and supply system propulsively vents through the enjoy at a controlled flowrate. During the long coast portion of the orbit, the LH2 tank is locked up (unvented). Thus, all of the vented H2 flow is used in the engine for thrust and none is wastefully vented overboard. The key to managing the tank pressure and therefore the H2 flow to the engine is to manage and balance the energy flow into the LH2 tank with the MLI and tank heaters with the energy flow out of the LH2 tank through the vented H2 flow. A moderate scale (71 cu ft) LH2 storage and supply system was installed and insulated at the NASA MSFC Test Area 300. The operation of the system is described in this paper. The test program for the LH2 system consisted of two parts: 1) a series of engineering tests to characterize the performance of the various components in the system: and 2) a 30-day simulation of a complete LEO and GEO transfer mission. This paper describes the results of the engineering tests, and correlates these results with analytical models used to design future advanced Solar Orbit Transfer Vehicles.

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.

High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner For Advanced Rocket Engines

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.; Ellis, David; Singh, Jogender

2014-01-01

Advanced high thermal conductivity materials research conducted at NASA Marshall Space Flight Center (MSFC) with state of the art combustion chamber liner material NARloy-Z showed that its thermal conductivity can be increased significantly by adding diamond particles and sintering it at high temperatures. For instance, NARloy-Z containing 40 vol. percent diamond particles, sintered at 975C to full density by using the Field assisted Sintering Technology (FAST) showed 69 percent higher thermal conductivity than baseline NARloy-Z. Furthermore, NARloy-Z-40vol. percent D is 30 percent lighter than NARloy-Z and hence the density normalized thermal conductivity is 140 percent better. These attributes will improve the performance and life of the advanced rocket engines significantly. By one estimate, increased thermal conductivity will directly translate into increased turbopump power up to 2X and increased chamber pressure for improved thrust and ISP, resulting in an expected 20 percent improvement in engine performance. Follow on research is now being conducted to demonstrate the benefits of this high thermal conductivity NARloy-Z-D composite for combustion chamber liner applications in advanced rocket engines. The work consists of a) Optimizing the chemistry and heat treatment for NARloy-Z-D composite, b) Developing design properties (thermal and mechanical) for the optimized NARloy-Z-D, c) Fabrication of net shape subscale combustion chamber liner, and d) Hot fire testing of the liner for performance. FAST is used for consolidating and sintering NARlo-Z-D. The subscale cylindrical liner with built in channels for coolant flow is also fabricated near net shape using the FAST process. The liner will be assembled into a test rig and hot fire tested in the MSFC test facility to determine performance. This paper describes the development of this novel high thermal conductivity NARloy-Z-D composite material, and the advanced net shape technology to fabricate the combustion chamber liner. Properties of optimized NARloy-Z-D composite material will also be presented.

Nuclear rocket using indigenous Martian fuel NIMF

NASA Technical Reports Server (NTRS)

Zubrin, Robert

1991-01-01

In the 1960's, Nuclear Thermal Rocket (NTR) engines were developed and ground tested capable of yielding isp of up to 900 s at thrusts up to 250 klb. Numerous trade studies have shown that such traditional hydrogen fueled NTR engines can reduce the inertial mass low earth orbit (IMLEO) of lunar missions by 35 percent and Mars missions by 50 to 65 percent. The same personnel and facilities used to revive the hydrogen NTR can also be used to develop NTR engines capable of using indigenous Martian volatiles as propellant. By putting this capacity of the NTR to work in a Mars descent/acent vehicle, the Nuclear rocket using Indigenous Martian Fuel (NIMF) can greatly reduce the IMLEO of a manned Mars mission, while giving the mission unlimited planetwide mobility.

Solar thermal power systems point-focusing thermal and electric applications projects. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Marriott, A.

1980-01-01

The activities of the Point-Focusing Thermal and Electric Applications (PETEA) project for the fiscal year 1979 are summarized. The main thrust of the PFTEA Project, the small community solar thermal power experiment, was completed. Concept definition studies included a small central receiver approach, a point-focusing distributed receiver system with central power generation, and a point-focusing distributed receiver concept with distributed power generation. The first experiment in the Isolated Application Series was initiated. Planning for the third engineering experiment series, which addresses the industrial market sector, was also initiated. In addition to the experiment-related activities, several contracts to industry were let and studies were conducted to explore the market potential for point-focusing distributed receiver (PFDR) systems. System analysis studies were completed that looked at PFDR technology relative to other small power system technology candidates for the utility market sector.

Plasma Engines,

DTIC Science & Technology

1982-09-08

low thrust, long duration power device, the plasma engine 6 has certain distinct advantages. For a chemical fuel rocket engine , a thrust of M.’)1...PLASMA ENGINES.CU) UNCLASSZICD FTO-ZIftS)T-0636-98 NL * UUUUU UUMile ~ FTD-ID(RS)T-0636-82 FOREIGN TECHNOLOGY DIVISION q 14 PLASMA ENGINES bv Sung...8 September 1982 MICROFICHE NR: FTD-82-C-001198 PLASMA ENGINES By: Sung Yuyang English pages: 7 Source: Hangkong Zhishi, March 1982, pp. 12-13 Country

Flight evaluation of an extended engine life mode on an F-15 airplane

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Conners, Timothy R.

1992-01-01

An integrated flight and propulsion control system designed to reduce the rate of engine deterioration was developed and evaluated in flight on the NASA Dryden F-15 research aircraft. The extended engine life mode increases engine pressure ratio while reducing engine airflow to lower the turbine temperature at constant thrust. The engine pressure ratio uptrim is modulated in real time based on airplane maneuver requirements, flight conditions, and engine information. The extended engine life mode logic performed well, significantly reducing turbine operating temperature. Reductions in fan turbine inlet temperature of up to 80 F were obtained at intermediate power and up to 170 F at maximum augmented power with no appreciable loss in thrust. A secondary benefit was the considerable reduction in thrust-specific fuel consumption. The success of the extended engine life mode is one example of the advantages gained from integrating aircraft flight and propulsion control systems.

Saturn Apollo Program

NASA Image and Video Library

1965-03-01

This photograph shows F-1 engines being stored in the F-1 Engine Preparation Shop, building 4666, at the Marshall Space Flight Center. Each F-1 engine produced a thrust of 1,500,000 pounds. A cluster of five engines was mounted on the thrust structure of the S-IC stage of a 364-foot long Saturn V launch vehicle that ultimately took astronauts to the Moon.

n/a

NASA Image and Video Library

1961-01-01

The static firing of a Saturn F-1 engine at the Marshall Space Flight Center's Static Test Stand. The F-1 engine is a single-start, 1,5000,000 Lb fixed-thrust, bipropellant rocket system. The engine uses liquid oxygen as the oxidizer and RP-1 (kerosene) as fuel. The five-engine cluster used on the first stage of the Saturn V produces 7,500,000 lbs of thrust.

Dynamics of high-bypass-engine thrust reversal using a variable-pitch fan

NASA Technical Reports Server (NTRS)

Schaefer, J. W.; Sagerser, D. R.; Stakolich, E. G.

1977-01-01

The test program demonstrated that successful and rapid forward-to reverse-thrust transients can be performed without any significant engine operational limitations for fan blade pitch changes through either feather pitch or flat pitch. For through-feather-pitch operation with a flight inlet, fan stall problems were encountered, and a fan blade overshoot technique was used to establish reverse thrust.

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.

Cruise Missile Engines

NASA Technical Reports Server (NTRS)

1982-01-01

Williams International's F107 fanjet engine is used in two types of cruise missiles, Navy-sponsored Tomahawk and the Air Force AGM-86B Air Launched Cruise Missile (ALCM). Engine produces about 600 pounds thrust, is one foot in diameter and weighs only 141 pounds. Design was aided by use of a COSMIC program in calculating airflows in engine's internal ducting, resulting in a more efficient engine with increased thrust and reduced fuel consumption.

A Layman's Guide to Thrust Engine Development for Super/Hyper Sonic Flight.

ERIC Educational Resources Information Center

Thiesse, James L.

The intention of this paper is to discuss the advances in thrust engines from the initial development of the J58/SR-71 (JT11D-20) of the U.S. Air Force's SR-71 Blackbird to the development of the RAM and SCRAM engines necessary to propel the new generations of high-flying super-speed aircraft. Engineering complexities suggest that the engines and…

Net thrust calculation sensitivity of an afterburning turbofan engine to variations in input parameters

NASA Technical Reports Server (NTRS)

Hughes, D. L.; Ray, R. J.; Walton, J. T.

1985-01-01

The calculated value of net thrust of an aircraft powered by a General Electric F404-GE-400 afterburning turbofan engine was evaluated for its sensitivity to various input parameters. The effects of a 1.0-percent change in each input parameter on the calculated value of net thrust with two calculation methods are compared. This paper presents the results of these comparisons and also gives the estimated accuracy of the overall net thrust calculation as determined from the influence coefficients and estimated parameter measurement accuracies.

Physical Limitations of Nuclear Propulsion for Earth to Orbit

NASA Technical Reports Server (NTRS)

Blevins, John A.; Patton, Bruce; Rhys, Noah O.; Schafer, Charles F. (Technical Monitor)

2001-01-01

An assessment of current nuclear propulsion technology for application in Earth to Orbit (ETO) missions has been performed. It can be shown that current nuclear thermal rocket motors are not sufficient to provide single stage performance as has been stated by previous studies. Further, when taking a systems level approach, it can be shown that NTRs do not compete well with chemical engines where thrust to weight ratios of greater than I are necessary, except possibly for the hybrid chemical/nuclear LANTR (LOX Augmented Nuclear Thermal Rocket) engine. Also, the ETO mission requires high power reactors and consequently large shielding weights compared to NTR space missions where shadow shielding can be used. In the assessment, a quick look at the conceptual ASPEN vehicle proposed in 1962 in provided. Optimistic NTR designs are considered in the assessment as well as discussion on other conceptual nuclear propulsion systems that have been proposed for ETO. Also, a quick look at the turbulent, convective heat transfer relationships that restrict the exchange of nuclear energy to thermal energy in the working fluid and consequently drive the reactor mass is included.

A Computational Study to Investigate the Effect of Altitude on Deteriorated Engine Performance

NASA Astrophysics Data System (ADS)

Koh, W. C.; Mazlan, N. M.; Rajendran, P.; Ismail, M. A.

2018-05-01

This study presents an investigation on the effect of operational altitudes on the performance of the deteriorated engine. A two-spool high bypass ratio turbofan engine is used as the test subject for this study. The engine is modelled in Gas Turbine Simulation Program (GSP) based on an existing engine model from literature. Real flight data were used for the validation. Deterioration rate of 0.1% per day is applied for all turbofan components engine. The simulation is performed by varying the altitude from sea level until 9000m. Results obtained show reduction in air mass flow rate and engine thrust as altitude increases. The reduction in air mass flow rate is due to the lower air density at higher altitude hence reduces amount of engine thrust. At 1000m to 4000m, thrust specific fuel consumption (TSFC) of the engine is improved compared to sea level. However depleted in TSFC is shown when the aircraft flies at altitude higher than 4000m. At this altitude, the effect of air density is dominant. As a result, the engine is required to burn more fuel to provide a higher thrust to sustain the aircraft speed. More fuel is consumed hence depletion in TSFC is obtained.

NACA Conference on Turbojet-Engine Thrust Augmentation Research: A Compilation of the Papers Presented by NACA Staff Members

NASA Technical Reports Server (NTRS)

1948-01-01

The conference on Turbojet-Engine Thrust-Augmentation Research was organized by the NACA to present in summarized form the results of the latest experimental and analytical investigations conducted at the Lewis Flight Propulsion Laboratory on methods of augmenting the thrust of turbojet engines. The technical discussions are reproduced herewith in the same form in which they were presented. The original presentation in this record are considered as complementary to, rather than substitutes for, the committee's system of complete and formal reports.

Performance Increase Verification for a Bipropellant Rocket Engine

NASA Technical Reports Server (NTRS)

Alexander, Leslie; Chapman, Jack; Wilson, Reed; Krismer, David; Lu, Frank; Wilson, Kim; Miller, Scott; England, Chris

2008-01-01

Component performance assessment testing for a, pressure-fed earth storable bipropellant rocket engine was successfully completed at Aerojet's Redmond test facility. The primary goal of the this development project is to increase the specific impulse of an apogee class bi-propellant engine to greater than 330 seconds with nitrogen tetroxide and monomethylhydrazine propellants and greater than 335 seconds with nitrogen tetroxide and hydrazine. The secondary goal of the project is to take greater advantage of the high temperature capabilities of iridium/rhenium chambers. In order to achieve these goals, the propellant feed pressures were increased to 400 psia, nominal, which in turn increased the chamber pressure and temperature, allowing for higher c*. The tests article used a 24-on-24 unlike doublet injector design coupled with a copper heat sink chamber to simulate a flight configuration combustion chamber. The injector is designed to produce a nominal 200 lbf of thrust with a specific impulse of 335 seconds (using hydrazine fuel). Effect of Chamber length on engine C* performance was evaluated with the use of modular, bolt-together test hardware and removable chamber inserts. Multiple short duration firings were performed to characterize injector performance across a range of thrust levels, 180 to 220 lbf, and mixture ratios, from 1.1 to 1.3. During firing, ignition transient, chamber pressure, and various temperatures were measured in order to evaluate the performance of the engine and characterize the thermal conditions. The tests successfully demonstrated the stable operation and performance potential of a full scale engine with a measured c* of XXXX ft/sec (XXXX m/s) under nominal operational conditions.

Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

2016-01-01

NARloy-Z alloy (Cu-3 percent, Ag-0.5 percent, Zr) is a state of the art alloy currently used for fabricating rocket engine combustion chamber liners. Research conducted at NASA-MSFC and Penn State – Applied Research Laboratory has shown that thermal conductivity of NARloy-Z can be increased significantly by adding diamonds to form a composite (NARloy-Z-D). NARloy-Z-D is also lighter than NARloy-Z. These attributes make this advanced composite material an ideal candidate for fabricating combustion chamber liner for an advanced rocket engine. Increased thermal conductivity will directly translate into increased turbopump power and increased chamber pressure for improved thrust and specific impulse. This paper describes the process development for fabricating a subscale high thermal conductivity NARloy-Z-D combustion chamber liner using Field Assisted Sintering Technology (FAST). The FAST process uses a mixture of NARloy-Z and diamond powders which is sintered under pressure at elevated temperatures. Several challenges were encountered, i.e., segregation of diamonds, machining the super hard NARloy-Z-D composite, net shape fabrication and nondestructive examination. The paper describes how these challenges were addressed. Diamonds coated with copper (CuD) appear to give the best results. A near net shape subscale combustion chamber liner is being fabricated by diffusion bonding cylindrical rings of NARloy-Z-CuD using the FAST process.

Advances in Thrust-Based Emergency Control of an Airplane

NASA Technical Reports Server (NTRS)

Creech, Gray; Burken, John J.; Burcham, Bill

2003-01-01

Engineers at NASA's Dryden Flight Research Center have received a patent on an emergency flight-control method implemented by a propulsion-controlled aircraft (PCA) system. Utilizing the preexisting auto-throttle and engine-pressure-ratio trim controls of the airplane, the PCA system provides pitch and roll control for landing an airplane safely without using aerodynamic control surfaces that have ceased to function because of a primary-flight-control-system failure. The installation of the PCA does not entail any changes in pre-existing engine hardware or software. [Aspects of the method and system at previous stages of development were reported in Thrust-Control System for Emergency Control of an Airplane (DRC-96-07), NASA Tech Briefs, Vol. 25, No. 3 (March 2001), page 68 and Emergency Landing Using Thrust Control and Shift of Weight (DRC-96-55), NASA Tech Briefs, Vol. 26, No. 5 (May 2002), page 58.]. Aircraft flight-control systems are designed with extensive redundancy to ensure low probabilities of failure. During recent years, however, several airplanes have exhibited major flight-control-system failures, leaving engine thrust as the last mode of flight control. In some of these emergency situations, engine thrusts were successfully modulated by the pilots to maintain flight paths or pitch angles, but in other situations, lateral control was also needed. In the majority of such control-system failures, crashes resulted and over 1,200 people died. The challenge lay in creating a means of sufficient degree of thrust-modulation control to safely fly and land a stricken airplane. A thrust-modulation control system designed for this purpose was flight-tested in a PCA an MD-11 airplane. The results of the flight test showed that without any operational control surfaces, a pilot can land a crippled airplane (U.S. Patent 5,330,131). The installation of the original PCA system entailed modifications not only of the flight-control computer (FCC) of the airplane but also of each engine-control computer. Inasmuch as engine-manufacturer warranties do not apply to modified engines, the challenge became one of creating a PCA system that does not entail modifications of the engine computers.

Full scale technology demonstration of a modern counterrotating unducted fan engine concept. Engine test

NASA Technical Reports Server (NTRS)

1987-01-01

The Unducted Fan (UDF) engine is an innovative aircraft engine concept based on an ungeared, counterrotating, unducted, ultra-high-bypass turbofan configuration. This engine is being developed to provide a high thrust-to-weight ratio power plant with exceptional fuel efficiency for subsonic aircraft application. This report covers the successful ground testing of this engine. A test program exceeding 100-hr duration was completed, in which all the major goals were achieved. The following accomplishments were demonstrated: (1) full thrust (25,000 lb); (2) full counterrotating rotor speeds (1393+ rpm); (3) low specific fuel consumption (less than 0.24 lb/hr/lb); (4) new composite fan design; (5) counterrotation of structures, turbines, and fan blades; (6) control system; (7) actuation system; and (8) reverse thrust.

The 727 airplane target thrust reverser static performance model test for refanned JT8D engines

NASA Technical Reports Server (NTRS)

Chow, C. T. P.; Atkey, E. N.

1974-01-01

The results of a scale model static performance test of target thrust reverser configurations for the Pratt and Whitney Aircraft JT8D-100 series engine are presented. The objective of the test was to select a series of suitable candidate reverser configurations for the subsequent airplane model wind tunnel ingestion and flight controls tests. Test results indicate that adequate reverse thrust performance with compatible engine airflow match is achievable for the selected configurations. Tapering of the lips results in loss of performance and only minimal flow directivity. Door pressure surveys were conducted on a selected number of lip and fence configurations to obtain data to support the design of the thrust reverser system.

Materials for Liquid Propulsion Systems. Chapter 12

NASA Technical Reports Server (NTRS)

Halchak, John A.; Cannon, James L.; Brown, Corey

2016-01-01

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

MD-11 PCA - View of aircraft on ramp

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 is taxiing to a position on the flightline at NASA's Dryden Flight Research Center, Edwards, California, following its completion of the first and second landings ever performed by a transport aircraft under engine power only (on Aug. 29, 1995). The milestone flight, with NASA research pilot and former astronaut Gordon Fullerton at the controls, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple. For pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

A transport aircraft lands for the first time under engine power only, as this McDonnell Douglas MD-11 touches down at 11:38 a.m., Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - Closeup view of aircraft on ramp

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 has taxied to a position on the flightline at NASA's Dryden Flight Research Center, Edwards, California, following its completion of the first and second landings ever performed by a transport aircraft under engine power only (on Aug. 29, 1995). The milestone flight, with NASA research pilot and former astronaut Gordon Fullerton at the controls, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple. For pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

A transport aircraft lands for the first time under engine power only, as this McDonnell Douglas MD-11 touches down at 11:38 a.m., Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Jr., Lee H. (Inventor)

1991-01-01

The invention is a method and system for monitoring and directly displaying the actual thrust produced by a jet aircraft engine under determined operating conditions and the available thrust and predicted (commanded) thrust of a functional model of an ideal engine under the same determined operating conditions. A first set of actual value output signals representative of a plurality of actual performance parameters of the engine under the determined operating conditions is generated and compared with a second set of predicted value output signals representative of the predicted value of corresponding performance parameters of a functional model of the engine under the determined operating conditions to produce a third set of difference value output signals within a range of normal, caution, or warning limit values. A thrust indicator displays when any one of the actual value output signals is in the warning range while shaping function means shape each of the respective difference output signals as each approaches the limit of the respective normal, caution, and warning range limits.

Space Storable Rocket Technology (SSRT) basic program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Mueller, T.; Casillas, A. R.; Huang, D.

1992-01-01

The Space Storable Rocket Technology Program (SSRT) was conducted to establish a technology for a new class of high performance and long life bipropellant engines using space storable propellants. The results are described. Task 1 evaluated several characteristics for a number of fuels to determine the best space storable fuel for use with LO2. The results indicated that LO2-N2H4 is the best propellant combination and provides the maximum mission/system capability maximum payload into GEO of satellites. Task 2 developed two models, performance and thermal. The performance model indicated the performance goal of specific impulse greater than or = 340 seconds (sigma = 204) could be achieved. The thermal model was developed and anchored to hot fire test data. Task 3 consisted of design, fabrication, and testing of a 200 lbf thrust test engine operating at a chamber pressure of 200 psia using LO2-N2H4. A total of 76 hot fire tests were conducted demonstrating performance greater than 340 (sigma = 204) which is a 25 second specific impulse improvement over the existing highest performance flight apogee type engines.

Innovative Double Bypass Engine for Increased Performance

NASA Astrophysics Data System (ADS)

Manoharan, Sanjivan

Engines continue to grow in size to meet the current thrust requirements of the civil aerospace industry. Large engines pose significant transportation problems and require them to be split in order to be shipped. Thus, large amounts of time have been spent in researching methods to increase thrust capabilities while maintaining a reasonable engine size. Unfortunately, much of this research has been focused on increasing the performance and efficiencies of individual components while limited research has been done on innovative engine configurations. This thesis focuses on an innovative engine configuration, the High Double Bypass Engine, aimed at increasing fuel efficiency and thrust while maintaining a competitive fan diameter and engine length. The 1-D analysis was done in Excel and then compared to the results from Numerical Propulsion Simulation System (NPSS) software and were found to be within 4% error. Flow performance characteristics were also determined and validated against their criteria.

Wind tunnel test of model target thrust reversers for the Pratt and Whitney aircraft JT8D-100 series engines installed on a 727-200 airplane

NASA Technical Reports Server (NTRS)

Hambly, D.

1974-01-01

The results of a low speed wind tunnel test of 0.046 scale model target thrust reversers installed on a 727-200 model airplane are presented. The full airplane model was mounted on a force balance, except for the nacelles and thrust reversers, which were independently mounted and isolated from it. The installation had the capability of simulating the inlet airflows and of supplying the correct proportions of primary and secondary air to the nozzles. The objectives of the test were to assess the compatibility of the thrust reversers target door design with the engine and airplane. The following measurements were made: hot gas ingestion at the nacelle inlets; model lift, drag, and pitching moment; hot gas impingement on the airplane structure; and qualitative assessment of the rudder effectiveness. The major parameters controlling hot gas ingestion were found to be thrust reverser orientation, engine power setting, and the lip height of the bottom thrust reverser doors on the side nacelles. The thrust reversers tended to increase the model lift, decrease the drag, and decrease the pitching moment.

A Regeneratively Cooled Thrust Chamber For The Fastrac Engine

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Sparks, Dave; Woodcock, Gordon

2000-01-01

Abstract This paper presents the development of a low-cost, regeneratively-cooled thrust chamber for the Fastrac engine. The chamber was fabricated using hydraformed copper tubing to form the coolant jacket and wrapped with a fiber reinforced polymer composite Material to form a structural jacket. The thrust chamber design and fabrication approach was based upon Space America. Inc.'s 12,000 lb regeneratively-cooled LOX/kerosene rocket engine. Fabrication of regeneratively cooled thrust chambers by tubewall construction dates back to the early US ballistic missile programs. The most significant innovations in this design was the development of a low-cost process for fabrication from copper tubing (nickel alloy was the usual practice) and use of graphite composite overwrap as the pressure containment, which yields an easily fabricated, lightweight pressure jacket around the copper tubes A regeneratively-cooled reusable thrust chamber can benefit the Fastrac engine program by allowing more efficient (cost and scheduler testing). A proof-of-concept test article has been fabricated and will he tested at Marshall Space Flight Center in the late Summer or Fall of 2000.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

NASA Astrophysics Data System (ADS)

Tian, Lu; Chen, Li-Hong; Chen, Qiang; Zhong, Feng-Quan; Chang, Xin-Yu

2016-02-01

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5-3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Characterization of advanced electric propulsion systems

NASA Technical Reports Server (NTRS)

Ray, P. K.

1982-01-01

Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

Emergency Flight Control Using Only Engine Thrust and Lateral Center-of-Gravity Offset: A First Look

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Burken, John; Maine, Trindel A.; Bull, John

1997-01-01

Normally, the damage that results in a total loss of the primary flight controls of a jet transport airplane, including all engines on one side, would be catastrophic. In response, NASA Dryden has conceived an emergency flight control system that uses only the thrust of a wing-mounted engine along with a lateral center-of-gravity (CGY) offset from fuel transfer. Initial analysis and simulation studies indicate that such a system works, and recent high-fidelity simulation tests on the MD-11 and B-747 suggest that the system provides enough control for a survivable landing. This paper discusses principles of flight control using only a wing engine thrust and CGY offset, along with the amount of CGY offset capability of some transport airplanes. The paper also presents simulation results of the throttle-only control capability and closed-loop control of ground track using computer-controlled thrust.

Orbit Transfer Vehicle (OTV) engine phase A study, extension 1. Volume 3: Study cost estimates

NASA Technical Reports Server (NTRS)

Christensen, K. L.

1980-01-01

Program cost and planning data based on 1980 technology and shown in 1979 dollars for a 20K lb Thrust Staged Combustion Cycle Engine are presented. These data were compared with those for the Advanced Expander Cycle Engine at 10K lb and 20K lb thrust levels.

4. COMPLETE X15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING ...

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

4. COMPLETE X-15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING STRUCTURE AT ENGINE END OF PLANE. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

Low thrust vehicle concept study

NASA Technical Reports Server (NTRS)

1980-01-01

Low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low acceleration orbit transfer of large space systems were defined. Results indicate that it is cost effective and least risk to combine the OTV and stowed spacecraft in a single 65 K Shuttle. The study shows that the engine for an optimized low thrust stage (1) does not require very low thrust; (2) 1-3 K thrust range appears optimum; (3) thrust transient is not a concern; (4) throttling probably not worthwhile; and (5) multiple thrusters complicate OTV/LSS design and aggravate LSS loads. Regarding the optimum vehicle for low acceleration missions, the single shuttle launch (LSS and expendable OTV) is most cost effective and least risky. Multiple shuttles increase diameter 20%. The space based radar structure short OTV (which maximizes space available for packaged LSS) favors use of torus tank. Propellant tank pressures/vapor residuals are little affected by engine thrust level or number of burns.

Analytical study of nozzle performance for nuclear thermal rockets

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.; Kacynski, Kenneth J.

1991-01-01

Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively.

Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

1995-01-01

The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.

Convective Heat Transfer with and without Film Cooling in High Temperature, Fuel Rich and Lean Environments

NASA Astrophysics Data System (ADS)

Greiner, Nathan J.

Modern turbine engines require high turbine inlet temperatures and pressures to maximize thermal efficiency. Increasing the turbine inlet temperature drives higher heat loads on the turbine surfaces. In addition, increasing pressure ratio increases the turbine coolant temperature such that the ability to remove heat decreases. As a result, highly effective external film cooling is required to reduce the heat transfer to turbine surfaces. Testing of film cooling on engine hardware at engine temperatures and pressures can be exceedingly difficult and expensive. Thus, modern studies of film cooling are often performed at near ambient conditions. However, these studies are missing an important aspect in their characterization of film cooling effectiveness. Namely, they do not model effect of thermal property variations that occur within the boundary and film cooling layers at engine conditions. Also, turbine surfaces can experience significant radiative heat transfer that is not trivial to estimate analytically. The present research first computationally examines the effect of large temperature variations on a turbulent boundary layer. Subsequently, a method to model the effect of large temperature variations within a turbulent boundary layer in an environment coupled with significant radiative heat transfer is proposed and experimentally validated. Next, a method to scale turbine cooling from ambient to engine conditions via non-dimensional matching is developed computationally and the experimentally validated at combustion temperatures. Increasing engine efficiency and thrust to weight ratio demands have driven increased combustor fuel-air ratios. Increased fuel-air ratios increase the possibility of unburned fuel species entering the turbine. Alternatively, advanced ultra-compact combustor designs have been proposed to decrease combustor length, increase thrust, or generate power for directed energy weapons. However, the ultra-compact combustor design requires a film cooled vane within the combustor. In both these environments, the unburned fuel in the core flow encounters the oxidizer rich film cooling stream, combusts, and can locally heat the turbine surface rather than the intended cooling of the surface. Accordingly, a method to quantify film cooling performance in a fuel rich environment is prescribed. Finally, a method to film cool in a fuel rich environment is experimentally demonstrated.

Performance deterioration based on simulated aerodynamic loads test, JT9D jet engine diagnostics program

NASA Technical Reports Server (NTRS)

Stromberg, W. J.

1981-01-01

An engine was specially prepared with extensive instrumentation to monitor performance, case temperatures, and clearance changes. A special loading device was used to apply known loads on the engine by the use of cables placed around the flight inlet. These loads simulated the estimated aerodynamic pressure distributions that occur on the inlet in various segments of a typical airplane flight. Test results indicate that the engine lost 1.3 percent in take-off thrust specific fuel consumption (TSFC) during the course of the test effort. Permanent clearance changes due to the loads accounted for 1.1 percent; increase in low pressure compressor airfoil roughness and thermal distortion in the high pressure turbine accounted for 0.2 percent. Pretest predicted performance loss due to clearance changes was 0.9 percent in TSFC. Therefore, the agreement between measurement and prediction is considered to be excellent.

Evaluation of Fuel Character Effects on J79 Engine Combustion System

DTIC Science & Technology

1979-06-01

A. Overall Engine Description The J79 engine is a lightweight, high-thrust, axial - flow turbojet engine with variable afterburner thrust. This engine...thimbles are arranged to provide flow patterns for flame stabilization in the primary zone and mixing and turbine inlet temperature profile control at...measured with stainard )S𔃾Z orifices- Fuel flow races uere measured with calibrated turbine flotaMcers corrected for the density aan viscosity of each

Direct and system effects of water ingestion into jet engine compresors

NASA Technical Reports Server (NTRS)

Murthy, S. N. B.; Ehresman, C. M.; Haykin, T.

1986-01-01

Water ingestion into aircraft-installed jet engines can arise both during take-off and flight through rain storms, resulting in engine operation with nearly saturated air-water droplet mixture flow. Each of the components of the engine and the system as a whole are affected by water ingestion, aero-thermally and mechanically. The greatest effects arise probably in turbo-machinery. Experimental and model-based results (of relevance to 'immediate' aerothermal changes) in compressors have been obtained to show the effects of film formation on material surfaces, centrifugal redistribution of water droplets, and interphase heat and mass transfer. Changes in the compressor performance affect the operation of the other components including the control and hence the system. The effects on the engine as a whole are obtained through engine simulation with specified water ingestion. The interest is in thrust, specific fuel consumption, surge margin and rotational speeds. Finally two significant aspects of performance changes, scalability and controllability, are discussed in terms of characteristic scales and functional relations.

Space rocket engine on the base of the reactor-pumped laser for the interplanetary flights and earth orbital applications

NASA Astrophysics Data System (ADS)

Gulevich, Andrey V.; Dyachenko, Peter P.; Kukharchuk, Oleg F.; Zrodnikov, Anatoly V.

2000-01-01

In this report the concept of vehicle-based reactor-laser engine for long time interplanetary and interorbital (LEO to GEO) flights is proposed. Reactor-pumped lasers offer the perspective way to create on the base of modern nuclear and lasers technologies the low mass and high energy density, repetitively pulsed vehicle-based laser of average power 100 kW. Nowadays the efficiency of nuclear-to-optical energy conversion reached the value of 2-3%. The demo model of reactor-pumped laser facility is under construction in Institute for Physics and Power Engineering (Obninsk, Russia). It enable us to hope that using high power laser on board of the vehicle could make the effective space laser engine possible. Such engine may provide the high specific impulse ~1000-2000 s with the thrust up to 10-100 n. Some calculation results of the characteristics of vehicle-based reactor-laser thermal engine concept are also presented. .

Test experience, 490 N high performance (321 sec Isp) engine

NASA Technical Reports Server (NTRS)

Schoenman, L.; Rosenberg, S. D.; Jassowski, D. M.

1992-01-01

Engines with area ratios of 44:1 and 286:1 are tested by means of hot fire tests using the NTO/MMH bipropellant to maximize the performance of the combined technologies. The low-thrust engine systems are designed with oxidation resistant materials that can operate at temperatures of more than 2204 C for tens of hours. The chamber is attached to the injector in a configuration that prevents overheating of the injector, valve, and the spacecraft interface. Three injectors with 44:1 area ratios are capable of nominal specific impulse values of 309 sec, and a performance of 321 lbf-sec/lbm is noted for an all-welded engine assembly with area ratio of 286:1. The all-welded engine is shown to have an acceptable design margin for thermal characteristics. High-performance liquid apogee engines are shown to perform optimally when based on iridium/rhenium chamber technology, use of a special platelet injector, and the minimization of losses due to fuel-film cooling.

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.

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.

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.

Rocket thrust chamber thermal barrier coatings

NASA Technical Reports Server (NTRS)

Quentmeyer, R. J.

1985-01-01

Subscale rocket thrust chamber tests were conducted to evaluate the effectiveness and durability of thin yttria stabilized zirconium oxide coatings applied to the thrust chamber hot-gas side wall. The fabrication consisted of arc plasma spraying the ceramic coating and bond coat onto a mandrell and then electrodepositing the copper thrust chamber wall around the coating. Chambers were fabricated with coatings .008, and .005 and .003 inches thick. The chambers were thermally cycled at a chamber pressure of 600 psia using oxygen-hydrogen as propellants and liquid hydrogen as the coolant. The thicker coatings tended to delaminate, early in the cyclic testing, down to a uniform sublayer which remained well adhered during the remaining cycles. Two chambers with .003 inch coatings were subjected to 1500 thermal cycles with no coating loss in the throat region, which represents a tenfold increase in life over identical chambers having no coatings. An analysis is presented which shows that the heat lost to the coolant due to the coating, in a rocket thrust chamber design having a coating only in the throat region, can be recovered by adding only one inch to the combustion chamber length.

Operability of an Ejector Enhanced Pulse Combustor in a Gas Turbine Environment

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Dougherty, Kevin

2008-01-01

A pressure-gain combustor comprised of a mechanically valved, liquid fueled pulsejet, an ejector, and an enclosing shroud, was coupled to a small automotive turbocharger to form a self-aspirating, thrust producing gas turbine engine. The system was constructed in order to investigate issues associated with the interaction of pulsed combustion devices and turbomachinery. Installed instrumentation allowed for sensing of distributed low frequency pressure and temperature, high frequency pressure in the shroud, fuel flow rate, rotational speed, thrust, and laboratory noise. The engine ran successfully and reliably, achieving a sustained thrust of 5 to 6 lbf, and maintaining a rotor speed of approximately 90,000 rpm, with a combustor pressure gain of approximately 4 percent. Numerical simulations of the system without pressure-gain combustion indicated that the turbocharger would not operate. Thus, the new combustor represented a substantial improvement in system performance. Acoustic measurements in the shroud and laboratory indicated turbine stage sound pressure level attenuation of 20 dB. This is consistent with published results from detonative combustion experiments. As expected, the mechanical reed valves suffered considerable damage under the higher pressure and thermal loading characteristics of this system. This result underscores the need for development of more robust valve systems for this application. The efficiency of the turbomachinery components did not appear to be significantly affected by unsteadiness associated with pulsed combustion, though the steady component efficiencies were already low, and thus not expected to be particularly sensitive.

75 FR 14377 - Airworthiness Directives; Pratt & Whitney (PW) Model PW2037, PW2037(M), and PW2040 Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-25

...The FAA proposes to adopt a new airworthiness directive (AD) for PW Model PW2037, PW2037(M), and PW2040 turbofan engines. This proposed AD would require removing erosion damage on fan blades with cutback leading edges and restoring the leading edge contour. This proposed AD results from reports from PW that fan blade leading edge erosion can result in a fan thrust deterioration mode (FTDM) condition, which reduces the engine's capability of producing full rated take-off thrust. We are proposing this AD to prevent loss of engine thrust from an FTDM condition, which could result in an inability to maintain safe flight.

Performance (Off-Design) Cycle Analysis for a Turbofan Engine With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This report presents the performance of a steady-state, dual-spool, separate-exhaust turbofan engine, with an interstage turbine burner (ITB) serving as a secondary combustor. The ITB, which is located in the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet-engine propulsion. A detailed off-design performance analysis of ITB engines is written in Microsoft(Registered Trademark) Excel (Redmond, Washington) macrocode with Visual Basic Application to calculate engine performances over the entire operating envelope. Several design-point engine cases are pre-selected using a parametric cycle-analysis code developed previously in Microsoft(Registered Trademark) Excel, for off-design analysis. The off-design code calculates engine performances (i.e. thrust and thrust-specific-fuel-consumption) at various flight conditions and throttle settings.

Performance Cycle Analysis of a Two-Spool, Separate-Exhaust Turbofan With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This paper presents the performance cycle analysis of a dual-spool, separate-exhaust turbofan engine, with an Interstage Turbine Burner serving as a secondary combustor. The ITB, which is located at the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet engine propulsion. A detailed performance analysis of this engine has been conducted for steady-state engine performance prediction. A code is written and is capable of predicting engine performances (i.e., thrust and thrust specific fuel consumption) at varying flight conditions and throttle settings. Two design-point engines were studied to reveal trends in performance at both full and partial throttle operations. A mission analysis is also presented to assure the advantage of saving fuel by adding ITB.

Performance of Simple Gas Foil Thrust Bearings in Air

NASA Technical Reports Server (NTRS)

Bruckner, Robert J.

2012-01-01

Foil bearings are self-acting hydrodynamics devices used to support high speed rotating machinery. The advantages that they offer to process fluid lubricated machines include: high rotational speed capability, no auxiliary lubrication system, non-contacting high speed operation, and improved damping as compared to rigid hydrodynamic bearings. NASA has had a sporadic research program in this technology for almost 6 decades. Advances in the technology and understanding of foil journal bearings have enabled several new commercial products in recent years. These products include oil-free turbochargers for both heavy trucks and automobiles, high speed electric motors, microturbines for distributed power generation, and turbojet engines. However, the foil thrust bearing has not received a complimentary level of research and therefore has become the weak link of oil-free turbomachinery. In an effort to both provide machine designers with basic performance parameters and to elucidate the underlying physics of foil thrust bearings, NASA Glenn Research Center has completed an effort to experimentally measure the performance of simple gas foil thrust bearing in air. The database includes simple bump foil supported thrust bearings with full geometry and manufacturing techniques available to the user. Test conditions consist of air at ambient pressure and temperatures up to 500 C and rotational speeds to 55,000 rpm. A complete set of axial load, frictional torque, and rotational speed is presented for two different compliant sub-structures and inter-pad gaps. Data obtained from commercially available foil thrust bearings both with and without active cooling is presented for comparison. A significant observation made possible by this data set is the speed-load capacity characteristic of foil thrust bearings. Whereas for the foil journal bearing the load capacity increases linearly with rotational speed, the foil thrust bearing operates in the hydrodynamic high speed limit. In this case, the load capacity is constant and in fact often decreases with speed if other factors such as thermal conditions and runner distortions are permitted to dominate the bearing performance.

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2012 CFR

2012-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2014 CFR

2014-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2010 CFR

2010-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2013 CFR

2013-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2011 CFR

2011-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

Preliminary flight evaluation of an engine performance optimization algorithm

NASA Technical Reports Server (NTRS)

Lambert, H. H.; Gilyard, G. B.; Chisholm, J. D.; Kerr, L. J.

1991-01-01

A performance seeking control (PSC) algorithm has undergone initial flight test evaluation in subsonic operation of a PW 1128 engined F-15. This algorithm is designed to optimize the quasi-steady performance of an engine for three primary modes: (1) minimum fuel consumption; (2) minimum fan turbine inlet temperature (FTIT); and (3) maximum thrust. The flight test results have verified a thrust specific fuel consumption reduction of 1 pct., up to 100 R decreases in FTIT, and increases of as much as 12 pct. in maximum thrust. PSC technology promises to be of value in next generation tactical and transport aircraft.

Design issues for lunar in situ aluminum/oxygen propellant rocket engines

NASA Technical Reports Server (NTRS)

Meyer, Michael L.

1992-01-01

Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

Dynamic Imbalance Would Counter Offcenter Thrust

NASA Technical Reports Server (NTRS)

Mccanna, Jason

1994-01-01

Dynamic imbalance generated by offcenter thrust on rotating body eliminated by shifting some of mass of body to generate opposing dynamic imbalance. Technique proposed originally for spacecraft including massive crew module connected via long, lightweight intermediate structure to massive engine module, such that artificial gravitation in crew module generated by rotating spacecraft around axis parallel to thrust generated by engine. Also applicable to dynamic balancing of rotating terrestrial equipment to which offcenter forces applied.

Direct Fusion Drive for a Human Mars Orbital Mission

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

Paluszek, Michael; Pajer, Gary; Razin, Yosef

2014-08-01

The Direct Fusion Drive (DFD) is a nuclear fusion engine that produces both thrust and electric power. It employs a field reversed configuration with an odd-parity rotating magnetic field heating system to heat the plasma to fusion temperatures. The engine uses deuterium and helium-3 as fuel and additional deuterium that is heated in the scrape-off layer for thrust augmentation. In this way variable exhaust velocity and thrust is obtained.

Using Engine Thrust for Emergency Flight Control: MD-11 and B-747 Results

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Maine, Trindel A.; Burken, John J.; Bull, John

1998-01-01

With modern digital control systems, using engine thrust for emergency flight control to supplement or replace failed aircraft normal flight controls has become a practical consideration. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control. An F-15 and an MD-11 airplane have been landed without using any flight control surfaces. Preliminary studies have also been conducted that show that engines on only one wing can provide some flight control capability if the lateral center of gravity can be shifted toward the side of the airplane that has the operating engine(s). Simulator tests of several airplanes with no flight control surfaces operating and all engines out on the left wing have all shown positive control capability within the available range of lateral center-of-gravity offset. Propulsion-controlled aircraft systems that can operate without modifications to engine control systems, thus allowing PCA technology to be installed on less capable airplanes or at low cost, are also desirable. Further studies have examined simplified 'PCA Lite' and 'PCA Ultralite' concepts in which thrust control is provided by existing systems such as auto-throttles or a combination of existing systems and manual pilot control.

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.

The impact of the fuel chemical composition on volatile organic compounds emitted by an in-service aircraft gas turbine engine

NASA Astrophysics Data System (ADS)

Setyan, A.; Kuo, Y. Y.; Brem, B.; Durdina, L.; Gerecke, A. C.; Heeb, N. V.; Haag, R.; Wang, J.

2017-12-01

Aircraft emissions received increased attention recently because of the steady growth of aviation transport in the last decades. Aircraft engines substantially contribute to emissions of particulate matter and gaseous pollutants in the upper and lower troposphere. Among all the pollutants emitted by aircrafts, volatile organic compounds (VOCs) are particularly important because they are mainly emitted at ground level, posing a serious health risk for people living or working near airports. A series of measurements was performed at the aircraft engine testing facility of SR Technics (Zürich airport, Switzerland). Exhausts from an in-service turbofan engine were sampled at the engine exit plane by a multi-point sampling probe. A wide range of instruments was connected to the common sampling line to determine physico-chemical characteristics of non-volatile particulate matter and gaseous pollutants. Conventional Jet A-1 fuel was used as the base fuel, and measurements were performed with the base fuel doped with two different mixtures of aromatic compounds (Solvesso 150 and naphthalene-depleted Solvesso 150) and an alternative fuel (hydro-processed esters and fatty acids [HEFA] jet fuel). During this presentation, we will show results obtained for VOCs. These compounds were sampled with 3 different adsorbing cartridges, and analyzed by thermal desorption gas chromatography/mass spectrometry (TD-GC/MS, for Tenax TA and Carboxen 569) and by ultra-performance liquid chromatography/ mass spectrometry (UPLC/MS, for DNPH). The total VOC concentration was also measured with a flame ionization detector (FID). In addition, fuel samples were also analyzed by GC/MS, and their chemical compositions were compared to the VOCs emitted via engine exhaust. Total VOCs concentrations were highest at ground idle (>200 ppm C at 4-7% thrust), and substantially lower at high thrust (<3 ppm C during take-off, 100% thrust). Fuel samples were dominated by alkanes, whereas VOCs emitted by the aircraft engine were mainly constituted of alkanes, oxygenated compounds, and aromatics. More than 50 % of the compounds identified in the exhaust were not present in the fuel, and thus were formed during combustion. The impact of the fuel doping with aromatics and the alternative fuel on VOCs emitted by the engine will also be discussed.

Investigation of electroforming techniques, literature analysis report

NASA Technical Reports Server (NTRS)

Malone, G. A.

1975-01-01

A literature analysis is presented of reports, specifications, and documented experiences with the use of electroforming to produce copper and nickel structures for aerospace and other engineering applications. The literature period covered is from 1948 to 1974. Specific effort was made to correlate mechanical property data for the electrodeposited material with known electroforming solution compositions and operating conditions. From this survey, electrolytes are suggested for selection to electroform copper and nickel outer shells on regeneratively cooled thrust chamber liners, and other devices subject to thermal and pressure exposure, based on mechanical properties obtainable, performance under various thermal environments, and ease of process control for product reproducibility. Processes of potential value in obtaining sound bonds between electrodeposited copper and nickel and copper alloy substrates are also discussed.

Guidance, Navigation, and Control Considerations for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Mitchell, Doyce P.; Kim, Tony

2015-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation NTP system could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of a first generation NTP in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NTP project could also help enable high performance fission power systems and Nuclear Electric Propulsion (NEP). Guidance, navigation, and control of NTP may have some unique but manageable characteristics.

Airflow and thrust calibration of an F100 engine, S/N P680059, at selected flight conditions

NASA Technical Reports Server (NTRS)

Biesiadny, T. J.; Lee, D.; Rodriguez, J. R.

1978-01-01

An airflow and thrust calibration of an F100 engine, S/N P680059, was conducted to study airframe propulsion system integration losses in turbofan-powered high-performance aircraft. The tests were conducted with and without thrust augmentation for a variety of simulated flight conditions with emphasis on the transonic regime. The resulting corrected airflow data generalized into one curve with corrected fan speed while corrected gross thrust increased as simulated flight conditions increased. Overall agreement between measured data and computed results was 1 percent for corrected airflow and -1 1/2 percent for gross thrust. The results of an uncertainty analysis are presented for both parameters at each simulated flight condition.

Uncertainty of in-flight thrust determination

NASA Technical Reports Server (NTRS)

Abernethy, Robert B.; Adams, Gary R.; Steurer, John W.; Ascough, John C.; Baer-Riedhart, Jennifer L.; Balkcom, George H.; Biesiadny, Thomas

1986-01-01

Methods for estimating the measurement error or uncertainty of in-flight thrust determination in aircraft employing conventional turbofan/turbojet engines are reviewed. While the term 'in-flight thrust determination' is used synonymously with 'in-flight thrust measurement', in-flight thrust is not directly measured but is determined or calculated using mathematical modeling relationships between in-flight thrust and various direct measurements of physical quantities. The in-flight thrust determination process incorporates both ground testing and flight testing. The present text is divided into the following categories: measurement uncertainty methodoogy and in-flight thrust measurent processes.

Thrust Stand for Electric Propulsion Performance Evaluation

NASA Technical Reports Server (NTRS)

Markusic, T. E.; Jones, J. E.; Cox, M. D.

2004-01-01

An electric propulsion thrust stand capable of supporting thrusters with total mass of up to 125 kg and 1 mN to 1 N thrust levels has been developed and tested. The mechanical design features a conventional hanging pendulum arm attached to a balance mechanism that transforms horizontal motion into amplified vertical motion, with accommodation for variable displacement sensitivity. Unlike conventional hanging pendulum thrust stands, the deflection is independent of the length of the pendulum arm, and no reference structure is required at the end of the pendulum. Displacement is measured using a non-contact, optical linear gap displacement transducer. Mechanical oscillations are attenuated using a passive, eddy current damper. An on-board microprocessor-based level control system, which includes a two axis accelerometer and two linear-displacement stepper motors, continuously maintains the level of the balance mechanism - counteracting mechanical %era drift during thruster testing. A thermal control system, which includes heat exchange panels, thermocouples, and a programmable recirculating water chiller, continuously adjusts to varying thermal loads to maintain the balance mechanism temperature, to counteract thermal drifts. An in-situ calibration rig allows for steady state calibration both prior to and during thruster testing. Thrust measurements were carried out on a well-characterized 1 kW Hall thruster; the thrust stand was shown to produce repeatable results consistent with previously published performance data.

MITEE-B: A Compact Ultra Lightweight Bi-Modal Nuclear Propulsion Engine for Robotic Planetary Science Missions

NASA Astrophysics Data System (ADS)

Powell, James; Maise, George; Paniagua, John; Borowski, Stanley

2003-01-01

Nuclear thermal propulsion (NTP) enables unique new robotic planetary science missions that are impossible with chemical or nuclear electric propulsion systems. A compact and ultra lightweight bi-modal nuclear engine, termed MITEE-B (MInature ReacTor EnginE - Bi-Modal) can deliver 1000's of kilograms of propulsive thrust when it operates in the NTP mode, and many kilowatts of continuous electric power when it operates in the electric generation mode. The high propulsive thrust NTP mode enables spacecraft to land and takeoff from the surface of a planet or moon, to hop to multiple widely separated sites on the surface, and virtually unlimited flight in planetary atmospheres. The continuous electric generation mode enables a spacecraft to replenish its propellant by processing in-situ resources, provide power for controls, instruments, and communications while in space and on the surface, and operate electric propulsion units. Six examples of unique and important missions enabled by the MITEE-B engine are described, including: (1) Pluto lander and sample return; (2) Europa lander and ocean explorer; (3) Mars Hopper; (4) Jupiter atmospheric flyer; (5) SunBurn hypervelocity spacecraft; and (6) He3 mining from Uranus. Many additional important missions are enabled by MITEE-B. A strong technology base for MITEE-B already exists. With a vigorous development program, it could be ready for initial robotic science and exploration missions by 2010 AD. Potential mission benefits include much shorter in-space times, reduced IMLEO requirements, and replenishment of supplies from in-situ resources.

Preliminary supersonic flight test evaluation of performance seeking control

NASA Technical Reports Server (NTRS)

Orme, John S.; Gilyard, Glenn B.

1993-01-01

Digital flight and engine control, powerful onboard computers, and sophisticated controls techniques may improve aircraft performance by maximizing fuel efficiency, maximizing thrust, and extending engine life. An adaptive performance seeking control system for optimizing the quasi-steady state performance of an F-15 aircraft was developed and flight tested. This system has three optimization modes: minimum fuel, maximum thrust, and minimum fan turbine inlet temperature. Tests of the minimum fuel and fan turbine inlet temperature modes were performed at a constant thrust. Supersonic single-engine flight tests of the three modes were conducted using varied after burning power settings. At supersonic conditions, the performance seeking control law optimizes the integrated airframe, inlet, and engine. At subsonic conditions, only the engine is optimized. Supersonic flight tests showed improvements in thrust of 9 percent, increases in fuel savings of 8 percent, and reductions of up to 85 deg R in turbine temperatures for all three modes. The supersonic performance seeking control structure is described and preliminary results of supersonic performance seeking control tests are given. These findings have implications for improving performance of civilian and military aircraft.

Speciation and chemical evolution of nitrogen oxides in aircraft exhaust near airports.

PubMed

Wood, Ezra C; Herndon, Scott C; Timko, Michael T; Yelvington, Paul E; Miake-Lye, Richard C

2008-03-15

Measurements of nitrogen oxides from a variety of commercial aircraft engines as part of the JETS-APEX2 and APEX3 campaigns show that NOx (NOx [triple bond] NO + NO2) is emitted primarily in the form of NO2 at idle thrust and NO at high thrust. A chemical kinetics combustion model reproduces the observed NO2 and NOx trends with engine power and sheds light on the relevant chemical mechanisms. Experimental evidence is presented of rapid conversion of NO to NO2 in the exhaust plume from engines at low thrust. The rapid conversion and the high NO2/NOx emission ratios observed are unrelated to ozone chemistry. NO2 emissions from a CFM56-3B1 engine account for approximately 25% of the NOx emitted below 3000 feet (916 m) and 50% of NOx emitted below 500 feet (153 m) during a standard ICAO (International Civil Aviation Organization) landing-takeoff cycle. Nitrous acid (HONO) accounts for 0.5% to 7% of NOy emissions from aircraft exhaust depending on thrust and engine type. Implications for photochemistry near airports resulting from aircraft emissions are discussed.

Feasibility Study of a Pressure-fed Engine for a Water Recoverable Space Shuttle Booster

NASA Technical Reports Server (NTRS)

Gerstl, E.

1972-01-01

Detailed mass properties are presented for a gimbaled, fixed thrust, regeneratively cooled engine having a coaxial pintle injector. The baseline design parameters for this engine are tabulated. Mass properties are also summarized for several other engine configurations i.e., a hinge nozzle using a Techroll seal, a gimbaled duct cooled engine and a regeneratively cooled engine using liquid injection thrust vector control (LITVC). Detailed engine analysis and design trade studies leading to the selection of a regeneratively cooled gimbaled engine and pertaining to the selection of the baseline design configuration are also given.

75 FR 8471 - Airworthiness Directives; Airbus Model A340-200 and A340-300 Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-02-25

... condition as: A review of A340 missions has demonstrated that CFM56-5C forward engine mount thrust links... thrust link would reduce the forward engine mounts' structural integrity and could eventually lead to... products. The MCAI states: A review of A340 missions has demonstrated that CFM56-5C forward engine mount...

J85 Rejuvenation Through Technology Insertion

DTIC Science & Technology

2000-10-01

and Sabre 75 business addition to military production, the J85 was jets . Number Model Produced Aircraft Type(s) Engine Type Thrust (lbs) J85-GE-4 740...REJUVENATION THROUGH TECHNOLOGY INSERTION T.A. Brisken, P.N. Howell, A.C. Ewing Military Engines Operation GE Aircraft Engines 1 Neumann Way Cincinnati...OH 45215, USA Summary thrust to weight ratio turbojet engines with potential application to early cruise missiles and drones. The history of the

Orbital Transfer Rocket Engine Technology. Advanced Engine Study, Task D.6 Final Report

DTIC Science & Technology

1992-06-01

PROPERTIES _- -,mr m" , MANUAL a PAQ *E,- 7.3.2.1.2. IA .A.2 ,C -- 70-t’ i Rl I _ N -’.±v-j-. .......-441I 0.2% YS Design Allowable • -’Moo 0 2W0" 6W...Storage External Radiation Environment ( Buried Engine) The engine thrust chamber would be cold to the touch even at full thrust operation from the

A new generation of high performance engines for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Rosenberg, Sanders D.; Schoenman, Leonard

1991-01-01

Experimental data validating advanced engine designs at three thrust levels (5, 15, and 100 lbF) is presented. All of the three engine designs considered employ a Moog bipropellant torque motor valve, platelet injector design, and iridium-lined rhenium combustion chamber. Attention is focused on the performance, robustness, duration, and flexibility characteristics of the engines. It is noted that the 5- and 15-lbF thrust engines can deliver a steady state specific impulse in excess of 310 lbF-sec/lbm at an area ratio of 150:1, while the 150-lbF thrust engines deliver a steady state specific impulse of 320 lbF-sec/lbm at an area ratio of 250:1. The hot-fire test results reveal specific impulse improvements of 15 to 25 sec over conventional fuel film cooled columbium chamber designs while operating at maximum chamber temperatures.

State variable modeling of the integrated engine and aircraft dynamics

NASA Astrophysics Data System (ADS)

Rotaru, Constantin; Sprinţu, Iuliana

2014-12-01

This study explores the dynamic characteristics of the combined aircraft-engine system, based on the general theory of the state variables for linear and nonlinear systems, with details leading first to the separate formulation of the longitudinal and the lateral directional state variable models, followed by the merging of the aircraft and engine models into a single state variable model. The linearized equations were expressed in a matrix form and the engine dynamics was included in terms of variation of thrust following a deflection of the throttle. The linear model of the shaft dynamics for a two-spool jet engine was derived by extending the one-spool model. The results include the discussion of the thrust effect upon the aircraft response when the thrust force associated with the engine has a sizable moment arm with respect to the aircraft center of gravity for creating a compensating moment.

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

NASA Technical Reports Server (NTRS)

Landry, K.

2005-01-01

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

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the... this subpart must be run with the reverser installed. In complying with this section, the power control... regimes of control operations are incorporated necessitating scheduling of the power-control lever motion...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2010 CFR

2010-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the... this subpart must be run with the reverser installed. In complying with this section, the power control... regimes of control operations are incorporated necessitating scheduling of the power-control lever motion...

Engineering research, development and technology report

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

Langland, R T

1999-02-01

Nineteen ninety-eight has been a transition year for Engineering, as we have moved from our traditional focus on thrust areas to a more focused approach with research centers. These five new centers of excellence collectively comprise Engineering's Science and Technology program. This publication summarizes our formative year under this new structure. Let me start by talking about the differences between a thrust area and a research center. The thrust area is more informal, combining an important technology with programmatic priorities. In contrast, a research center is directly linked to an Engineering core technology. It is the purer model, for itmore » is more enduring yet has the scope to be able to adapt quickly to evolving programmatic priorities. To put it another way, the mission of a thrust area was often to grow the programs in conjunction with a technology, whereas the task of a research center is to vigorously grow our core technologies. By cultivating each core technology, we in turn enable long-term growth of new programs.« less

Energy Efficient Engine program advanced turbofan nacelle definition study

NASA Technical Reports Server (NTRS)

Howe, David C.; Wynosky, T. A.

1985-01-01

Advanced, low drag, nacelle configurations were defined for some of the more promising propulsion systems identified in the earlier Benefit/Cost Study, to assess the benefits associated with these advanced technology nacelles and formulate programs for developing these nacelles and low volume thrust reversers/spoilers to a state of technology readiness in the early 1990's. The study results established the design feasibility of advanced technology, slim line nacelles applicable to advanced technology, high bypass ratio turbofan engines. Design feasibility was also established for two low volume thrust reverse/spoiler concepts that meet or exceed the required effectiveness for these engines. These nacelle and thrust reverse/spoiler designs were shown to be applicable in engines with takeoff thrust sizes ranging from 24,000 to 60,000 pounds. The reduced weight, drag, and cost of the advanced technology nacelle installations relative to current technology nacelles offer a mission fuel burn savings ranging from 3.0 to 4.5 percent and direct operating cost plus interest improvements from 1.6 to 2.2 percent.

Selected Performance Measurements of the F-15 Active Axisymmetric Thrust-vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1998-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Selected Performance Measurements of the F-15 ACTIVE Axisymmetric Thrust-Vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1999-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Orbital transfer rocket engine technology 7.5K-LB thrust rocket engine preliminary design

NASA Technical Reports Server (NTRS)

Harmon, T. J.; Roschak, E.

1993-01-01

A preliminary design of an advanced LOX/LH2 expander cycle rocket engine producing 7,500 lbf thrust for Orbital Transfer vehicle missions was completed. Engine system, component and turbomachinery analysis at both on design and off design conditions were completed. The preliminary design analysis results showed engine requirements and performance goals were met. Computer models are described and model outputs are presented. Engine system assembly layouts, component layouts and valve and control system analysis are presented. Major design technologies were identified and remaining issues and concerns were listed.

Nuclear-Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rom, Frank E.

1968-01-01

The three basic types of nuclear power-plants (solid, liquid, and gas core) are compared on the bases of performance potential and the status of current technology. The solid-core systems are expected to have impulses in the range of 850 seconds, any thrust level (as long as it is greater than 10,000 pounds (44,480 newtons)), and thrust-to-engine-weight ratios of 2 to 20 pounds per pound (19.7 to 197 newtons per kilogram). There is negligible or no fuel loss from the solid-core system. The solid-core system, of course, has had the most work done on it. Large-scale tests have been performed on a breadboard engine that has produced specific impulses greater than 700 seconds at thrust levels of about 50,000 pounds (222,000 newtons). The liquid-core reactor would be interesting in the specific impulse range of 1200 to 1500 seconds. Again, any thrust level can be obtained depending on how big or small the reactor is made. The thrust-to-engine weight ratio for these systems would be in the range of 1 to 10. The discouraging feature of the liquid-core system is the high fuel-loss ratio anticipated. Values of 0.01 to 0.1 pound (0.00454 to 0.0454 kilograms) or uranium loss per pound (0.454 kilograms) of hydrogen are expected, if impulses in the range of 1200 to 1500 seconds are desired. The gas-core reactor shows specific impulses in the range of 1500 to 2500 seconds. The thrust levels should be at least as high as the weight so that the thrust-to-weight ratio does not go below 1. Because the engine weight is not expected to be under 100,000 pounds (444,800 newtons), thrust levels higher than 100,000 pounds (448,000 newtons) are of interest. The thrust-to-engine weights, in that case, would run from 1 to 20 pounds per pound (9.8 to 19.7 kilograms). Gas-core reactors tend to be very large, and can have high thrust-to-weight ratios. As in the case of the liquid-core system, the fuel loss that will be attendant with gas cores as envisioned today will be rather high. The loss rates will be 0.01 to 0.1 pound of uranium (0.00454 to 0.0454 kilograms) for each pound (0.454 kilograms) of hydrogen.

Rapid deceleration mode evaluation

NASA Technical Reports Server (NTRS)

Conners, Timothy R.; Nobbs, Steven G.; Orme, John S.

1995-01-01

Aircraft with flight capability above 1.4 normally have an RPM lockup or similar feature to prevent inlet buzz that would occur at low engine airflows. This RPM lockup has the effect of holding the engine thrust level at the intermediate power (maximum non-afterburning). For aircraft such as military fighters or supersonic transports, the need exists to be able to rapidly slow from supersonic to subsonic speeds. For example, a supersonic transport that experiences a cabin decompression needs to be able to slow/descend rapidly, and this requirement may size the cabin environmental control system. For a fighter, there may be a desire to slow/descend rapidly, and while doing so to minimize fuel usage and engine exhaust temperature. Both of these needs can be aided by achieving the minimum possible overall net propulsive force. As the intermediate power thrust levels of engines increase, it becomes even more difficult to slow rapidly from supersonic speeds. Therefore, a mode of the performance seeking control (PSC) system to minimize overall propulsion system thrust has been developed and tested. The rapid deceleration mode reduces the engine airflow consistent with avoiding inlet buzz. The engine controls are trimmed to minimize the thrust produced by this reduced airflow, and moves the inlet geometry to degrade the inlet performance. As in the case of the other PSC modes, the best overall performance (in this case the least net propulsive force) requires an integrated optimization of inlet, engine, and nozzle variables. This paper presents the predicted and measured results for the supersonic minimum thrust mode, including the overall effects on aircraft deceleration.

Kadenancy effect, acoustical resonance effect valveless pulse jet engine

NASA Astrophysics Data System (ADS)

Ismail, Rafis Suizwan; Jailani, Azrol; Haron, Muhammad Adli

2017-09-01

A pulse jet engine is a tremendously simple device, as far as moving parts are concerned, that is capable of using a range of fuels, an ignition device, and the ambient air to run an open combustion cycle at rates commonly exceeding 100 Hz. The pulse jet engine was first recognized as a worthy device for aeronautics applications with the introduction of the German V-1 Rocket, also known as the "Buzz Bomb." Although pulse jets are somewhat inefficient compared to other jet engines in terms of fuel usage, they have an exceptional thrust to weight ratio if the proper materials are chosen for its construction. For this reason, many hobbyists have adopted pulse jet engines for a propulsive device in RC planes, go-karts, and other recreational applications. The concept behind the design and function of propulsion devices are greatly inspired by the Newton's second and third laws. These laws quantitatively described thrust as a reaction force. Basically, whenever a mass is accelerated or expelled from one direction by a system, such a mass will exert the same force which will be equal in magnitude, however that will be opposite in direction over the same system. Thrust is that force utilized over a facade in a direction normal and perpendicular to the facade which is known as the thrust. This is the simplest explanation of the concept, on which propulsion devices functions. In mechanical engineering, any force that is orthogonal to the main load is generally referred to as thrust [1].

Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

NASA Astrophysics Data System (ADS)

Sathiyanathan, Kartheephan

This thesis describes the design, fabrication, and testing of a solid propellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

Performance of 4600-pound-thrust centrifugal-flow-type turbojet engine with water-alcohol injection at inlet

NASA Technical Reports Server (NTRS)

Glasser, Philip W

1950-01-01

An experimental investigation of the effects of injecting a water-alcohol mixture of 2:1 at the compressor inlet of a centrifugal-flow type turbojet engine was conducted in an altitude test chamber at static sea-level conditions and at an altitude of 20,000 feet with a flight Mach number of 0.78 with an engine operating at rated speed. The net thrust was augmented by 0.16 for both flight conditions with a ratio of injected liquid to air flow of 0.05. Further increases in the liquid-air ratio did not give comparable increases in thrust.

Analysis of Material Sample Heated by Impinging Hot Hydrogen Jet in a Non-Nuclear Tester

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Foote, John; Litchford, Ron

2006-01-01

A computational conjugate heat transfer methodology was developed and anchored with data obtained from a hot-hydrogen jet heated, non-nuclear materials tester, as a first step towards developing an efficient and accurate multiphysics, thermo-fluid computational methodology to predict environments for hypothetical solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, thermally radiating, unstructured-grid, and pressure-based formulation. The multiphysics invoked in this study include hydrogen dissociation kinetics and thermodynamics, turbulent flow, convective and thermal radiative, and conjugate heat transfers. Predicted hot hydrogen jet and material surface temperatures were compared with those of measurement. Predicted solid temperatures were compared with those obtained with a standard heat transfer code. The interrogation of physics revealed that reactions of hydrogen dissociation and recombination are highly correlated with local temperature and are necessary for accurate prediction of the hot-hydrogen jet temperature.

Reverse Core Engine with Thrust Reverser

NASA Technical Reports Server (NTRS)

Chandler, Jesse M. (Inventor); Suciu, Gabriel L. (Inventor)

2017-01-01

An engine system has a gas generator, a bi-fi wall surrounding at least a portion of the gas generator, a casing surrounding a fan, and the casing having first and second thrust reverser doors which in a deployed position abut each other and the bi-fi wall.

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

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.

Saturn Apollo Program

NASA Image and Video Library

1962-06-07

This photograph depicts the Rocketdyne static firing of the F-1 engine at the towering 76-meter Test Stand 1-C in Area 1-125 of the Edwards Air Force Base in California. The Saturn V S-IC (first) stage utilized five F-1 engines for its thrust. Each engine provided 1,500,000 pounds, for a combined thrust of 7,500,000 pounds with liquid oxygen and kerosene as its propellants.

U.S. Air Force Bomber Sustainment and Modernization: Background and Issues for Congress

DTIC Science & Technology

2014-06-04

turbofan Thrust: Each engine up to 17,000 pounds Wingspan: 185 feet (56.4 meters) Length: 159 feet, 4 inches (48.5 meters) Height: 40 feet, 8...precision and non-precision weapons. Features The B-1B’s blended wing and body configuration, variable-geometry wings, and turbofan afterburning engines... turbofan engine with afterburner Thrust: 30,000-plus pounds with afterburner, per engine Wingspan: 137 feet (41.8 meters) extended forward, 79 feet

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2012 CFR

2012-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2011 CFR

2011-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2013 CFR

2013-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

Multiphysics Thermal-Fluid Design Analysis of a Non-Nuclear Tester for Hot-Hydrogen Materials and Component Development

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Foote, John; Litchford, Ron

2006-01-01

The objective of this effort is to perform design analyses for a non-nuclear hot-hydrogen materials tester, as a first step towards developing efficient and accurate multiphysics, thermo-fluid computational methodology to predict environments for hypothetical solid-core, nuclear thermal engine thrust chamber design and analysis. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, thermally radiating, unstructured-grid, and pressure-based formulation. The multiphysics invoked in this study include hydrogen dissociation kinetics and thermodynamics, turbulent flow, convective, and thermal radiative heat transfers. The goals of the design analyses are to maintain maximum hot-hydrogen jet impingement energy and to minimize chamber wall heating. The results of analyses on three test fixture configurations and the rationale for final selection are presented. The interrogation of physics revealed that reactions of hydrogen dissociation and recombination are highly correlated with local temperature and are necessary for accurate prediction of the hot-hydrogen jet temperature.

Nuclear Thermal Rocket - Arc Jet Integrated System Model

NASA Technical Reports Server (NTRS)

Taylor, Brian D.; Emrich, William

2016-01-01

In the post-shuttle era, space exploration is moving into a new regime. Commercial space flight is in development and is planned to take on much of the low earth orbit space flight missions. With the development of a heavy lift launch vehicle, the Space Launch, System, NASA has become focused on deep space exploration. Exploration into deep space has traditionally been done with robotic probes. More ambitious missions such as manned missions to asteroids and Mars will require significant technology development. Propulsion system performance is tied to the achievability of these missions and the requirements of other developing technologies that will be required. Nuclear thermal propulsion offers a significant improvement over chemical propulsion while still achieving high levels of thrust. Opportunities exist; however, to build upon what would be considered a standard nuclear thermal engine to attain improved performance, thus further enabling deep space missions. This paper discuss the modeling of a nuclear thermal system integrated with an arc jet to further augment performance. The performance predictions and systems impacts are discussed.

Advanced materials for radiation-cooled rockets

NASA Technical Reports Server (NTRS)

Reed, Brian; Biaglow, James; Schneider, Steven

1993-01-01

The most common material system currently used for low thrust, radiation-cooled rockets is a niobium alloy (C-103) with a fused silica coating (R-512A or R-512E) for oxidation protection. However, significant amounts of fuel film cooling are usually required to keep the material below its maximum operating temperature of 1370 C, degrading engine performance. Also the R-512 coating is subject to cracking and eventual spalling after repeated thermal cycling. A new class of high-temperature, oxidation-resistant materials are being developed for radiation-cooled rockets, with the thermal margin to reduce or eliminate fuel film cooling, while still exceeding the life of silicide-coated niobium. Rhenium coated with iridium is the most developed of these high-temperature materials. Efforts are on-going to develop 22 N, 62 N, and 440 N engines composed of these materials for apogee insertion, attitude control, and other functions. There is also a complimentary NASA and industry effort to determine the life limiting mechanisms and characterize the thermomechanical properties of these materials. Other material systems are also being studied which may offer more thermal margin and/or oxidation resistance, such as hafnium carbide/tantalum carbide matrix composites and ceramic oxide-coated iridium/rhenium chambers.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Lee H., Jr. (Inventor)

1988-01-01

The invention is believed a major improvement that will have a broad application in governmental and commercial aviation. It provides a dynamic method and system for monitoring and simultaneously displaying in easily scanned form the available, predicted, and actual thrust of a jet aircraft engine under actual operating conditions. The available and predicted thrusts are based on the performance of a functional model of the aircraft engine under the same operating conditions. Other critical performance parameters of the aircraft engine and functional model are generated and compared, the differences in value being simultaneously displayed in conjunction with the displayed thrust values. Thus, the displayed information permits the pilot to make power adjustments directly while keeping him aware of total performance at a glance of a single display panel.

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Advanced expander cycle engine optimization

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1979-01-01

The performance optimization of expander cycle engines at vacuum thrust levels of 10K, 15K, and 20K lb is discussed. The optimization is conducted for a maximum engine length with an extendible nozzle in the retracted position of 60 inches and an engine mixture ratio of 6.0:1. The thrust chamber geometry and cycle analyses are documented. In addition, the sensitivity of a recommended baseline expander cycle to component performance variations is determined and chilldown/start propellant consumptions are estimated.

Performance and Environmental Test Results of the High Voltage Hall Accelerator Engineering Development Unit

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Haag, Thomas; Huang, Wensheng; Shastry, Rohit; Pinero, Luis; Peterson, Todd; Mathers, Alex

2012-01-01

NASA Science Mission Directorate's In-Space Propulsion Technology Program is sponsoring the development of a 3.5 kW-class engineering development unit Hall thruster for implementation in NASA science and exploration missions. NASA Glenn and Aerojet are developing a high fidelity high voltage Hall accelerator that can achieve specific impulse magnitudes greater than 2,700 seconds and xenon throughput capability in excess of 300 kilograms. Performance, plume mappings, thermal characterization, and vibration tests of the high voltage Hall accelerator engineering development unit have been performed. Performance test results indicated that at 3.9 kW the thruster achieved a total thrust efficiency and specific impulse of 58%, and 2,700 sec, respectively. Thermal characterization tests indicated that the thruster component temperatures were within the prescribed material maximum operating temperature limits during full power thruster operation. Finally, thruster vibration tests indicated that the thruster survived the 3-axes qualification full-level random vibration test series. Pre and post-vibration test performance mappings indicated almost identical thruster performance. Finally, an update on the development progress of a power processing unit and a xenon feed system is provided.

Monopropellant hydrazine resistoject: Engineering model fabrication and test task

NASA Technical Reports Server (NTRS)

Murch, C. K.

1973-01-01

The monopropellant hydrazine resistojet, termed the electrothermal hydrazine thruster (EHT) by TRW systems, thermally decomposes anhydrous hydrazine propellant to produce a high-temperature, low-molecular-weight gas for expulsion through a propulsive nozzle. The EHT developed for this program required about 3-5 watts of electrical power and produced 0.020 to 0.070 pound of thrust over the inlet pressure range of 100 to 400 psia. The thruster was designed for both pulsed and steady state operation. A summary of the GSFC original requirements and GSFC modified requirements, and the performance of the engineering model EHT is given. The experimental program leading to the engineering model EHT design, modifications necessary to achieve the required thruster life capability, and the results of the life test prgram. Other facets of the program, including analyses, preliminary design, specifications, data correlation, and recommendations for a flight model are discussed.

Experimental Evaluation of Cermet Turbine Stator Blades for Use at Elevated Gas Temperatures

NASA Technical Reports Server (NTRS)

Chiarito, Patrick T.; Johnston, James R.

1959-01-01

The suitability of cermets for turbine stator blades of a modified turbojet engine was determined at an average turbine-inlet-gas temperature of 2000 F. Such an increase in temperature would yield a premium in thrust from a service engine. Because the cermet blades require no cooling, all the available compressor bleed air could be used to cool a turbine made from conventional ductile alloys. Cermet blades were first run in 100-hour endurance tests at normal gas temperatures in order to evaluate two methods for mounting them. The elevated gas-temperature test was then run using the method of support considered best for high-temperature operation. After 52 hours at 2000 F, one of the group of four cermet blades fractured probably because of end loads resulting from thermal distortion of the spacer band of the nozzle diaphragm. Improved design of a service engine would preclude this cause of premature failure.

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.

Acoustics and Trust of Separate-Flow Exhaust Nozzles With Mixing Devices for High-Bypass-Ratio Engines

NASA Technical Reports Server (NTRS)

Saiyed, Naseem H.; Mikkelsen, Kevin L.; Bridges, James E.

2000-01-01

The NASA Glenn Research Center recently completed an experimental study to reduce the jet noise from modern turbofan engines. The study concentrated on exhaust nozzle designs for high-bypass-ratio engines. These designs modified the core and fan nozzles individually and simultaneously. Several designs provided an ideal jet noise reduction of over 2.5 EPNdB for the effective perceived noise level (EPNL) metric. Noise data, after correcting for takeoff thrust losses, indicated over a 2.0-EPNdB reduction for nine designs. Individually modifying the fan nozzle did not provide attractive EPNL reductions. Designs in which only the core nozzle was modified provided greater EPNL reductions. Designs in which core and fan nozzles were modified simultaneously provided the greatest EPNL reduction. The best nozzle design had a 2.7-EPNdB reduction (corrected for takeoff thrust loss) with a 0.06-point cruise thrust loss. This design simultaneously employed chevrons on the core and fan nozzles. In comparison with chevrons, tabs appeared to be an inefficient method for reducing jet noise. Data trends indicate that the sum of the thrust losses from individually modifying core and fan nozzles did not generally equal the thrust loss from modifying them simultaneously. Flow blockage from tabs did not scale directly with cruise thrust loss and the interaction between fan flow and the core nozzle seemed to strongly affect noise and cruise performance. Finally, the nozzle configuration candidates for full-scale engine demonstrations are identified.

Effects of Altitude on Turbojet Engine Performance

NASA Technical Reports Server (NTRS)

Fleming, William A

1951-01-01

Component and over-all performance characteristics of several turbojet engines investigated in the altitude facilities of the NACA Lewis Laboratory during the last several years are summarized to indicate the effects of altitude on turbojet engine performance. Data presented show that failure of turbojet engine performance to generalize for all altitudes can be traced to reductions in compressor efficiency, corrected air flow, and combustion efficiency at altitude. In addition, it is shown that although engines of different design may have equal thrusts at sea level, the thrusts at altitude may vary widely because of differences in compressor performance characteristics from one engine to another.

Full Scale Technology Demonstration of a Modern Counterrotating Unducted Fan Engine Concept. Design Report

NASA Technical Reports Server (NTRS)

1987-01-01

The Unducted Fan engine (UDF trademark) concept is based on an ungeared, counterrotating, unducted, ultra-high-bypass turbofan configuration. This engine is being developed to provide a high thrust-to-weight ratio power plant with exceptional fuel efficiency for subsonic aircraft application. This report covers the design methodology and details for the major components of this engine. The design intent of the engine is to efficiently produce 25,000 pounds of static thrust while meeting life and stress requirements. The engine is required to operate at Mach numbers of 0.8 or above.

Analysis of Factors Affecting the Performance of RLV Thrust Cell Liners

NASA Technical Reports Server (NTRS)

Arnold, Steven M. (Technical Monitor); Butler, Daniel T., Jr.; Pinders, Marek-Jerzy

2004-01-01

The reusable launch vehicle (RLV) thrust cell liner, or thrust chamber, is a critical component of the Space Shuttle Main Engine (SSME). It is designed to operate in some of the most severe conditions seen in engineering practice. This requirement, in conjunction with experimentally observed 'dog-house' failure modes characterized by bulging and thinning of the cooling channel wall, provides the motivation to study the factors that influence RLV thrust cell liner performance. Factors or parameters believed to be directly related to the observed characteristic deformation modes leading to failure under in-service loading conditions are identified, and subsequently investigated using the cylindrical version of the higher-order theory for functionally graded materials in conjunction with the Robinson's unified viscoplasticity theory and the power-law creep model for modeling the response of the liner s constituents. Configurations are analyzed in which specific modifications in cooling channel wall thickness or constituent materials are made to determine the influence of these parameters on the deformations resulting in the observed failure modes in the outer walls of the cooling channel. The application of thermal barrier coatings and functional grading are also investigated within this context. Comparison of the higher-order theory results based on the Robinson and power-law creep model predictions has demonstrated that, using the available material parameters, the power-law creep model predicts more precisely the experimentally observed deformation leading to the 'dog-house' failure mode for multiple short cycles, while also providing much improved computational efficiency. However, for a single long cycle, both models predict virtually identical deformations. Increasing the power-law creep model coefficients produces appreciable deformations after just one long cycle that would normally be obtained after multiple cycles, thereby enhancing the efficiency of the analysis. This provides a basis for the development of an accelerated modeling procedure to further characterize dog-house deformation modes in RLV thrust cell liners. Additionally, the results presented herein have demonstrated that the mechanism responsible for deformation leading to 'dog-house' failure modes is driven by pressure, creep/relaxation and geometric effects.

Flight Measurements of the Effect of a Controllable Thrust Reverser on the Flight Characteristics of a Single-Engine Jet Airplane

NASA Technical Reports Server (NTRS)

Anderson, Seth B.; Cooper, George E.; Faye, Alan E., Jr.

1959-01-01

A flight investigation was undertaken to determine the effect of a fully controllable thrust reverser on the flight characteristics of a single-engine jet airplane. Tests were made using a cylindrical target-type reverser actuated by a hydraulic cylinder through a "beep-type" cockpit control mounted at the base of the throttle. The thrust reverser was evaluated as an in-flight decelerating device, as a flight path control and airspeed control in landing approach, and as a braking device during the ground roll. Full deflection of the reverser for one reverser configuration resulted in a reverse thrust ratio of as much as 85 percent, which at maximum engine power corresponded to a reversed thrust of 5100 pounds. Use of the reverser in landing approach made possible a wide selection of approach angles, a large reduction in approach speed at steep approach angles, improved control of flight path angle, and more accuracy in hitting a given touchdown point. The use of the reverser as a speed brake at lower airspeeds was compromised by a longitudinal trim change. At the lower airspeeds and higher engine powers there was insufficient elevator power to overcome the nose-down trim change at full reverser deflection.

Thrust Slip Rates as a Control on the Presence and Spatial Distribution of High Metamorphic Heating Rates in Collisional Systems: The "Hot Iron" Model Revisited

NASA Astrophysics Data System (ADS)

Thigpen, R.; Ashley, K. T.; Law, R. D.; Mako, C. A.

2017-12-01

In natural systems, two key observations indicate that major strain discontinuities such as faults and shear zones should play a fundamental role in orogenic thermal evolution: (1) Large faults and shear zones often separate components of the composite orogen that have experienced broadly different thermal and deformational histories, and (2) quantitative metamorphic and diffusional studies indicate that heating rates are much faster and the duration of peak conditions much shorter in natural collisional systems than those predicted by numerical continuum deformation models. Because heat transfer processes such as conduction usually operate at much slower time scales than rates of other tectonic processes, thermal evolution is often transient and thus can be strongly influenced by tectonic disturbances that occur at rates much faster than thermal relaxation. Here, we use coupled thermal-mechanical finite element models of thrust faults to explore how fault slip rate may fundamentally influence the thermal evolution of individual footwall and hanging wall thrust slices. The model geometry involves a single crustal-scale thrust with a dip of 25° that is translated up the ramp at average velocities of 20, 35, and 50 km Myr-1, interpreted to represent average to relatively high slip rates observed in many collisional systems. Boundary conditions include crustal radioactive heat production, basal mantle heat flow, and surface erosion rates that are a function of thrust rate and subsequent topography generation. In the models, translation of the hanging wall along the crustal-scale detachment results in erosion, exhumation, and retrograde metamorphism of the emerging hanging wall topography and coeval burial, `hot iron' heating, and prograde metamorphism of the thrust footwall. Thrust slip rates of 20, 35, and 50 km Myr-1 yield maximum footwall heating rates ranging from 55-90° C Myr-1 and maximum hanging wall cooling rates of 138-303° C Myr-1. These relatively rapid heating rates explain, in part, the presence of chemical diffusion profiles in metamorphic minerals that are indicative of high heating rates. Additionally, the relatively high cooling rates explain preservation of chemical zoning, as rapid cooling prevents diffusive profiles from being substantially modified during exhumation.

The Design and Testing of a Miniature Turbofan Engine

NASA Technical Reports Server (NTRS)

Cosentino, Gary B.; Murray, James E.

2009-01-01

Off-the-shelf jet propulsion in the 50 - 500 lb thrust class sparse. A true twin-spool turbofan in this range does not exist. Adapting an off-the-shelf turboshaft engine is feasible. However the approx.10 Hp SPT5 can t quite make 50 lbs. of thrust. Packaging and integration is challenging, especially the exhaust. Building on our engine using a 25 Hp turboshaft seems promising if the engine becomes available. Test techniques used, though low cost, adequate for the purpose.

Development of HIDEC adaptive engine control systems

NASA Technical Reports Server (NTRS)

Landy, R. J.; Yonke, W. A.; Stewart, J. F.

1986-01-01

The purpose of NASA's Highly Integrated Digital Electronic Control (HIDEC) flight research program is the development of integrated flight propulsion control modes, and the evaluation of their benefits aboard an F-15 test aircraft. HIDEC program phases are discussed, with attention to the Adaptive Engine Control System (ADECS I); this involves the upgrading of PW1128 engines for operation at higher engine pressure ratios and the production of greater thrust. ADECS II will involve the development of a constant thrust mode which will significantly reduce turbine operating temperatures.

Air Taxi at Your Service

NASA Technical Reports Server (NTRS)

2002-01-01

Under an exclusive agreement with Eclipse Aviation Corporation, Williams International is manufacturing the EJ22 engine, a commercial version of the NASA/Williams FJX-2, for the Eclipse 500 aircraft. The new engine, which weighs approximately 85 pounds and delivers over 770 pounds of thrust, provides a higher thrust-to-weight ratio than any commercial turbofan ever produced. Being the smallest, quietest, and lightest commercial aircraft engine currently available, the EJ22 engine makes a whole new class of twinjet light aircraft feasible.

Camera Layout Design for the Upper Stage Thrust Cone

NASA Technical Reports Server (NTRS)

Wooten, Tevin; Fowler, Bart

2010-01-01

Engineers in the Integrated Design and Analysis Division (EV30) use a variety of different tools to aid in the design and analysis of the Ares I vehicle. One primary tool in use is Pro-Engineer. Pro-Engineer is a computer-aided design (CAD) software that allows designers to create computer generated structural models of vehicle structures. For the Upper State thrust cone, Pro-Engineer was used to assist in the design of a layout for two camera housings. These cameras observe the separation between the first and second stage of the Ares I vehicle. For the Ares I-X, one standard speed camera was used. The Ares I design calls for two separate housings, three cameras, and a lighting system. With previous design concepts and verification strategies in mind, a new layout for the two camera design concept was developed with members of the EV32 team. With the new design, Pro-Engineer was used to draw the layout to observe how the two camera housings fit with the thrust cone assembly. Future analysis of the camera housing design will verify the stability and clearance of the camera with other hardware present on the thrust cone.

Electric propulsion system technology

NASA Technical Reports Server (NTRS)

Brophy, John R.; Garner, Charles E.; Goodfellow, Keith D.; Pivirotto, Thomas J.; Polk, James E.

1992-01-01

The work performed in fiscal year (FY) 1991 under the Propulsion Technology Program RTOP (Research and Technology Objectives and Plans) No. (55) 506-42-31 for Low-Thrust Primary and Auxiliary Propulsion technology development is described. The objectives of this work fall under two broad categories. The first of these deals with the development of ion engines for primary propulsion in support of solar system exploration. The second with the advancement of steady-state magnetoplasmadynamic (MPD) thruster technology at 100 kW to multimegawatt input power levels. The major technology issues for ion propulsion are demonstration of adequate engine life at the 5 to 10 kW power level and scaling ion engines to power levels of tens to hundreds of kilowatts. Tests of a new technique in which the decelerator grid of a three-grid ion accelerator system is biased negative of neutralizer common potential in order to collect facility induced charge-exchange ions are described. These tests indicate that this SAND (Screen, Accelerator, Negative Decelerator) configuration may enable long duration ion engine endurance tests to be performed at vacuum chamber pressures an order of magnitude higher than previously possible. The corresponding reduction in pumping speed requirements enables endurance tests of 10 kW class ion engines to be performed within the resources of existing technology programs. The results of a successful 5,000-hr endurance of a xenon hollow cathode operating at an emission current of 25 A are described, as well as the initial tests of hollow cathodes operating on a mixture of argon and 3 percent nitrogen. Work performed on the development of carbon/carbon grids, a multi-orifice hollow cathode, and discharge chamber erosion reduction through the addition of nitrogen are also described. Critical applied-field MPD thruster technical issues remain to be resolved, including demonstration of reliable steady-state operation at input powers of hundreds to thousands of kilowatts, achievement of thruster efficiency and specific impulse levels required for missions of interest, and demonstration of adequate engine life at these input power, efficiency, and specific impulse levels. To address these issues we have designed, built, and tested a 100 kW class, radiation-cooled applied-field MPD thruster and a unique dual-beam thrust stand that enables separate measurements of the applied- and self-field thrust components. We have also initiated the development of cathode thermal and plasma sheath models that will eventually be used to guide the experimental program. In conjunction with the cathode modeling, a new cathode test facility is being constructed. This facility will support the study of cathode thermal behavior and erosion mechanisms, the diagnosis of the near-cathode plasma and the development and endurance testing of new, high-current cathode designs. To facilitate understanding of electrode surface phenomenon, we have implemented a telephoto technique to obtain photographs of the electrodes during engine operation. In order to reduce the background vacuum tank pressure during steady-state engine operation in order to obtain high fidelity anode thermal data, we have developed and are evaluating a gas-dynamic diffuser. A review of experience with alkali metal propellants for MPD thrusters led to the conclusion that alkali metals, particularly lithium, offer the potential for significant engine performance and lifetime improvements. These propellants are also condensible at room temperature, substantially reducing test facility pumping requirements. The most significant systems-level issue is the potential for spacecraft contamination. Subsequent experimental and theoretical efforts should be directed toward verifying the performance and lifetime gains and characterizing the thruster flow field to assess its impact on spacecraft surfaces. Consequently, we have begun the design and development of a new facility to study engine operation with alkali metal propellants.

Simulator Evaluation of Simplified Propulsion-Only Emergency Flight Control Systems on Transport Aircraft

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Kaneshige, John; Bull, John; Maine, Trindel A.

1999-01-01

With the advent of digital engine control systems, considering the use of engine thrust for emergency flight control has become feasible. Many incidents have occurred in which engine thrust supplemented or replaced normal aircraft flight controls. In most of these cases, a crash has resulted, and more than 1100 lives have been lost. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control capability. Using this PCA system, an F-15 and an MD-11 airplane have been landed without using any flight controls. In simulations, C-17, B-757, and B-747 PCA systems have also been evaluated successfully. These tests used full-authority digital electronic control systems on the engines. Developing simpler PCA systems that can operate without full-authority engine control, thus allowing PCA technology to be installed on less capable airplanes or at lower cost, is also a desire. Studies have examined simplified ?PCA Ultralite? concepts in which thrust control is provided using an autothrottle system supplemented by manual differential throttle control. Some of these concepts have worked well. The PCA Ultralite study results are presented for simulation tests of MD-11, B-757, C-17, and B-747 aircraft.

Thrust stand evaluation of engine performance improvement algorithms in an F-15 airplane

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1992-01-01

Results are presented from the evaluation of the performance seeking control (PSC) optimization algorithm developed by Smith et al. (1990) for F-15 aircraft, which optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. Comparisons are presented between the load cell measurements, PSC onboard model thrust calculations, and posttest state variable model computations. Actual performance improvements using the PSC algorithm are presented for its various modes. The results of using PSC algorithm are compared with similar test case results using the HIDEC algorithm.

The Rolls Royce Allison RB580 turbofan - Matching the market requirement for regional transport

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

Sadler, J.H.R.; Peacock, N.J.; Snyder, L.

1989-01-01

The RB580 high bypass turbofan engine has a thrust growth capability to 10,000 lb and has been optimized for efficient operation in regional markets involving 50-70 seat airliners with higher-than-turboprop cruise speeds. The two-spool engine configuration achieves an overall pressure ratio of 24 and features a single-stage wide-chord fan for high efficiency/low noise operation. The highly modular design of the configuration facilitates maintenance and repair; a dual-redundant full-authority digital electronic control system is incorporated. An SFC reduction of the order of 10 percent at cruise thrust is achieved, relative to current engines of comparable thrust class.

Subsonic Performance of Ejector Systems

NASA Astrophysics Data System (ADS)

Weil, Samuel

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

Easy method of matching fighter engine to airframe for use in aircraft engine design courses

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

Mattingly, J.D.

1989-01-01

The proper match of the engine(s) to the airframe affects both aircraft size and life cycle cost. A fast and straightforward method is developed and used for the matching of fighter engine(s) to airframes during conceptual design. A thrust-lapse equation is developed for the dual-spool, mixed-flow, afterburning turbofan type of engine based on the installation losses of 'Aircraft Engine Design' and the performance predictions of the cycle analysis programs ONX and OFFX. Using system performance requirements, the effects of aircraft thrust-to-weight, wing loading, and engine cycle on takeoff weight are analyzed and example design course results presented. 5 refs.

Thrust Augmentation with Mixer/Ejector Systems

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

DC-9 Flight Demonstration Program with Refanned JT8D Engines. Volume 3; Performance and Analysis

NASA Technical Reports Server (NTRS)

1975-01-01

The JT8D-109 engine has a sea level static, standard day bare engine takeoff thrust of 73,840 N. At sea level standard day conditions the additional thrust of the JT8D-109 results in 2,040 kg additional takeoff gross weight capability for a given field length. Range loss of the DC-9 Refan airplane for long range cruise was determined. The Refan airplane demonstrated stall, static longitudinal stability, longitudinal control, longitudinal trim, minimum control speeds, and directional control characteristics similar to the DC-9-30 production airplane and complied with airworthiness requirements. Cruise, climb, and thrust reverser performance were evaluated. Structural and dynamic ground test, flight test and analytical results substantiate Refan Program requirements that the nacelle, thrust reverser hardware, and the airplane structural modifications are flightworthy and certifiable and that the airplane meets flutter speed margins. Estimated unit cost of a DC-9 Refan retrofit program is 1.338 million in mid-1975 dollars with about an equal split in cost between airframe and engine.

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 deg, respectively.

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 degrees, respectively.

Heat pipe technology for advanced rocket thrust chambers

NASA Technical Reports Server (NTRS)

Rousar, D. C.

1971-01-01

The application of heat pipe technology to the design of rocket engine thrust chambers is discussed. Subjects presented are: (1) evaporator wick development, (2) specific heat pipe designs and test results, (3) injector design, fabrication, and cold flow testing, and (4) preliminary thrust chamber design.

Thrust law effects on the long-period modes of aerospace craft

NASA Technical Reports Server (NTRS)

Markopoulos, Nikos; Mease, Kenneth D.; Vinh, Nguyen X.

1989-01-01

An analytical study is presented of the longitudinal long-period dynamics of an aerospace craft in a nearly circular orbit, with a thrust law depending arbitrarily on the speed and altitude. A plane of engine possibilities is first defined, with points corresponding to propulsion systems having prescribed thrust slopes with respect to speed and altitude. Approximate expressions for the characteristic roots and times are obtained by first identifying a small quantity in the coefficients of the characteristic equation, and then expanding in a perturbation series about the origin of the plane of engine possibilities, for which the solution is always known. These expressions agree very well with the exact solutions over a wide range of altitudes and thrust laws. The period of the oscillatory translational mode (phugoid) is found to be independent to first order of the thrust law, generalizing results found by previous investigators for specific thrust laws. The results apply to the speed range from hypersonic to orbital.

Improving the breed - Shuttle development

NASA Technical Reports Server (NTRS)

Brand, V.

1985-01-01

An evaluation is made of design improvements that have been made to the Space Shuttle System, and the performance gains obtained; the most important of these stem from efforts to refine procedures for rendezvous with stricken satellites, in order to repair them. Ascent performance has been improved through Space Shuttle Main Engine thrust improvements and external tank weight reductions. On-orbit living convenience has been enhanced by the addition of small sleeping compartments and a galley. Greater flexibility has been obtained for reentry and landing maneuvers. Attention is given to problems which continue to be posed by the thermal protection tiles.

Effects of an in-flight thrust reverser on the stability and control characteristics of a single-engine fighter airplane model

NASA Technical Reports Server (NTRS)

Mercer, C. E.; Maiden, D. L.

1972-01-01

The changes in thrust minus drag performance as well as longitudinal and directional stability and control characteristics of a single-engine jet aircraft attributable to an in-flight thrust reverser of the blocker-deflector door type were investigated in a 16-foot transonic wind tunnel. The longitudinal and directional stability data are presented. Test conditions simulated landing approach conditions as well as high speed maneuvering such as may be required for combat or steep descent from high altitude.

LEO-to-GEO low thrust chemical propulsion

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1980-01-01

One approach being considered for transporting large space structures from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) is the use of low thrust chemical propulsion systems. A variety of chemical rocket engine cycles evaluated for this application for oxygen/hydrogen and oxygen/hydrocarbon propellants (oxygen/methane and oxygen/RF-1) are discussed. These cycles include conventional propellant turbine drives, turboalternator/electric motor pump drive, and fuel cell/electric motor pump drive as well as pressure fed engines. Thrust chamber cooling analysis results are presented for regenerative/radiation and film/radiation cooling.

Aircraft emissions of methane and nitrous oxide during the alternative aviation fuel experiment.

PubMed

Santoni, Gregory W; Lee, Ben H; Wood, Ezra C; Herndon, Scott C; Miake-Lye, Richard C; Wofsy, Steven C; McManus, J Barry; Nelson, David D; Zahniser, Mark S

2011-08-15

Given the predicted growth of aviation and the recent developments of alternative aviation fuels, quantifying methane (CH(4)) and nitrous oxide (N(2)O) emission ratios for various aircraft engines and fuels can help constrain projected impacts of aviation on the Earth's radiative balance. Fuel-based emission indices for CH(4) and N(2)O were quantified from CFM56-2C1 engines aboard the NASA DC-8 aircraft during the first Alternative Aviation Fuel Experiment (AAFEX-I) in 2009. The measurements of JP-8 fuel combustion products indicate that at low thrust engine states (idle and taxi, or 4% and 7% maximum rated thrusts, respectively) the engines emit both CH(4) and N(2)O at a mean ± 1σ rate of 170 ± 160 mg CH(4) (kg Fuel)(-1) and 110 ± 50 mg N(2)O (kg Fuel)(-1), respectively. At higher thrust levels corresponding to greater fuel flow and higher engine temperatures, CH(4) concentrations in engine exhaust were lower than ambient concentrations. Average emission indices for JP-8 fuel combusted at engine thrusts between 30% and 100% of maximum rating were -54 ± 33 mg CH(4) (kg Fuel)(-1) and 32 ± 18 mg N(2)O (kg Fuel)(-1), where the negative sign indicates consumption of atmospheric CH(4) in the engine. Emission factors for the synthetic Fischer-Tropsch fuels were statistically indistinguishable from those for JP-8.

High Thrust-to-Power Annular Engine Technology

NASA Technical Reports Server (NTRS)

Patterson, Michael J.; Thomas, Robert E.; Crofton, Mark W.; Young, Jason A.; Foster, John E.

2015-01-01

Gridded ion engines have the highest efficiency and total impulse of any mature electric propulsion technology, and have been successfully implemented for primary propulsion in both geocentric and heliocentric environments with excellent ground/in-space correlation of performance. However, they have not been optimized to maximize thrust-to-power, an important parameter for Earth orbit transfer applications. This publication discusses technology development work intended to maximize this parameter. These activities include investigating the capabilities of a non-conventional design approach, the annular engine, which has the potential of exceeding the thrust-to-power of other EP technologies. This publication discusses the status of this work, including the fabrication and initial tests of a large-area annular engine. This work is being conducted in collaboration among NASA Glenn Research Center, The Aerospace Corporation, and the University of Michigan.

High Thrust-to-Power Annular Engine Technology

NASA Technical Reports Server (NTRS)

Patterson, Michael; Thomas, Robert; Crofton, Mark; Young, Jason A.; Foster, John E.

2015-01-01

Gridded ion engines have the highest efficiency and total impulse of any mature electric propulsion technology, and have been successfully implemented for primary propulsion in both geocentric and heliocentric environments with excellent ground-in-space correlation of performance. However, they have not been optimized to maximize thrust-to-power, an important parameter for Earth orbit transfer applications. This publication discusses technology development work intended to maximize this parameter. These activities include investigating the capabilities of a non-conventional design approach, the annular engine, which has the potential of exceeding the thrust-to-power of other EP technologies. This publication discusses the status of this work, including the fabrication and initial tests of a large-area annular engine. This work is being conducted in collaboration among NASA Glenn Research Center, The Aerospace Corporation, and the University of Michigan.

In-flight thrust determination on a real-time basis

NASA Technical Reports Server (NTRS)

Ray, R. J.; Carpenter, T.; Sandlin, T.

1984-01-01

A real time computer program was implemented on a F-15 jet fighter to monitor in-flight engine performance of a Digital Electronic Engine Controlled (DEES) F-100 engine. The application of two gas generator methods to calculate in-flight thrust real time is described. A comparison was made between the actual results and those predicted by an engine model simulation. The percent difference between the two methods was compared to the predicted uncertainty based on instrumentation and model uncertainty and agreed closely with the results found during altitude facility testing. Data was obtained from acceleration runs of various altitudes at maximum power settings with and without afterburner. Real time in-flight thrust measurement was a major advancement to flight test productivity and was accomplished with no loss in accuracy over previous post flight methods.

Fuel-Efficient Descent and Landing Guidance Logic for a Safe Lunar Touchdown

NASA Technical Reports Server (NTRS)

Lee, Allan Y.

2011-01-01

The landing of a crewed lunar lander on the surface of the Moon will be the climax of any Moon mission. At touchdown, the landing mechanism must absorb the load imparted on the lander due to the vertical component of the lander's touchdown velocity. Also, a large horizontal velocity must be avoided because it could cause the lander to tip over, risking the life of the crew. To be conservative, the worst-case lander's touchdown velocity is always assumed in designing the landing mechanism, making it very heavy. Fuel-optimal guidance algorithms for soft planetary landing have been studied extensively. In most of these studies, the lander is constrained to touchdown with zero velocity. With bounds imposed on the magnitude of the engine thrust, the optimal control solutions typically have a "bang-bang" thrust profile: the thrust magnitude "bangs" instantaneously between its maximum and minimum magnitudes. But the descent engine might not be able to throttle between its extremes instantaneously. There is also a concern about the acceptability of "bang-bang" control to the crew. In our study, the optimal control of a lander is formulated with a cost function that penalizes both the touchdown velocity and the fuel cost of the descent engine. In this formulation, there is not a requirement to achieve a zero touchdown velocity. Only a touchdown velocity that is consistent with the capability of the landing gear design is required. Also, since the nominal throttle level for the terminal descent sub-phase is well below the peak engine thrust, no bound on the engine thrust is used in our formulated problem. Instead of bangbang type solution, the optimal thrust generated is a continuous function of time. With this formulation, we can easily derive analytical expressions for the optimal thrust vector, touchdown velocity components, and other system variables. These expressions provide insights into the "physics" of the optimal landing and terminal descent maneuver. These insights could help engineers to achieve a better "balance" between the conflicting needs of achieving a safe touchdown velocity, a low-weight landing mechanism, low engine fuel cost, and other design goals. In comparing the computed optimal control results with the preflight landing trajectory design of the Apollo-11 mission, we noted interesting similarities between the two missions.

Test Results for a Non-toxic, Dual Thrust Reaction Control Engine

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.; Turpin, Alicia A.

2005-01-01

A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized LOX/Ethanol propellants; and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber pressure (Pc) from 0.78 to 5.96 psia, and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing tests explored EPW from 0.080 to 10 seconds and DC from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215 pulses, and primary testing accumulated a total of 2,060 seconds of firing time and 3,646 pulses.

Investigation of two-dimensional wedge exhaust nozzles for advanced aircraft

NASA Technical Reports Server (NTRS)

Maiden, D. L.; Petit, J. E.

1975-01-01

Two-dimensional wedge nozzle performance characteristics were investigated in a series of wind-tunnel tests. An isolated single-engine/nozzle model was used to study the effects of internal expansion area ratio, aftbody cowl boattail angle, and wedge length. An integrated twin-engine/nozzle model, tested with and without empenage surfaces, included cruise, acceleration, thrust vectoring and thrust reversing nozzle operating modes. Results indicate that the thrust-minus-aftbody drag performance of the twin two-dimensional nozzle integration is significantly higher, for speeds greater than Mach 0.8, than the performance achieved with twin axisymmetric nozzle installations. Significant jet-induced lift was obtained on an aft-mounted lifting surface using a cambered wedge center body to vector thrust. The thrust reversing capabilities of reverser panels installed on the two-dimensional wedge center body were very effective for static or in-flight operation.

Orbit Transfer Rocket Engine Technology - 7.5K-LB Thrust Rocket Engine Preliminary Design

DTIC Science & Technology

1993-10-15

AND SPACE ADMINISTRATION October, 1993 r W NASA-Lewis Research Center Cleveland, Ohio 44135 94-08572 Contract Nc. NAS3-23773 Task B.7 and D.5 4I3’OA4 3 ...APPROACH 1 4.0 SUMMARY OF ACCOMPLISHMENTS 2 5.0 TECHNICAL DISCUSSIONS 3 6.0 PROGRAM WORK PLAN 5 6.1 Engine Analysis 5 6.2 Component Analysis 15 6.2.1...FIGURES Page Figure 1 Advanced Engine Studv Logic Diagram 4 Figure 2 Design Point Engine Pertormance at Full Thrust & MR = 6.0 7 Figure 3 Off-Design

Some effects of thermal-cycle-induced deformation in rocket thrust chambers

NASA Technical Reports Server (NTRS)

Hannum, N. P.; Price, R. G., Jr.

1981-01-01

The deformation process observed in the hot gas side wall of rocket combustion chambers was investigaged for three different liner materials. Five thrust chambers were cycled to failure by using hydrogen and oxygen as propellants at a chamber pressure of 4.14 MN/cu m. The deformation was observed nondestructively at midlife points and destructively after failure occurred. The cyclic life results are presented with an accompanying discussion about the problems of life prediction associated with the types of failures encountered in the present work. Data indicating the deformation of the thrust chamber liner as cycles are accumulated are presented for each of the test thrust chambers. From these deformation data and observation of the failure sites it is evident that modeling the failure process as classic low cycle thermal fatigue is inadequate as a life prediction method.

Development of circulation control technology for powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1987-01-01

The flow entraining capabilities of the Circulation Control Wing high lift system were employed to provide an even stronger STOL potential when synergistically combined with upper surface mounted engines. The resulting configurations generate very high supercirculation lift in addition to a vertical component of the pneumatically deflected engine thrust. A series of small scale wind tunnel tests and full scale static thrust deflection tests are discussed which provide a sufficient data base performance. These tests results show thrust deflections of greater than 90 deg produced pneumatically by nonmoving aerodynamic surfaces, and the ability to maintain constant high lift while varying the propulsive force from high thrust recovery required for short takeoff to high drag generation required for short low speed landings.

Closeup view of the aft fuselage of the Orbiter Discovery ...

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

Close-up view of the aft fuselage of the Orbiter Discovery looking at the thrust structure that supports the Space Shuttle Main Engines (SSMEs). In this view, SSME number two position is on the left and SSME number three position is on the right. The thrust structure transfers the forces produce by the engines into and through the airframe of the orbiter. The thrust structure includes the SSMEs load reaction truss structure, engine interface fittings and the hydraulic-actuator support structure. The propellant feed lines are the plugged and capped orifices within the engine bays. Note that SSME position two is rotated ninety degrees from position three and one. This was needed to enable enough clearance for the engines to fit and gimbal. Note in engine bay three is a clear view of the actuators that control the gambling of that engine. This view was taken from a service platform in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Deimos Methane-Oxygen Rocket Engine Test Results

NASA Astrophysics Data System (ADS)

Engelen, S.; Souverein, L. J.; Twigt, D. J.

This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

Development Status of the NASA MC-1 (Fastrac) Engine

NASA Technical Reports Server (NTRS)

Ballard, Richard O.; Olive, Tim; Turner, James E. (Technical Monitor)

2000-01-01

The MC-1 (formerly known as the Fastrac 60K) Engine is being developed for the X-34 technology demonstrator vehicle. It is a pump-fed liquid rocket engine with fixed thrust operating at one rated power level of 60,000 lbf vacuum thrust using a 15:1 area ratio nozzle (slightly higher for the 30:1 flight nozzle). Engine system development testing of the MC-1 has been ongoing since 24 Oct 1998. To date, 48 tests have been conducted on three engines using three separate test stands. This paper will provide some details of the engine, the tests conducted, and the lessons learned to date.

Model-Based Control of an Aircraft Engine using an Optimal Tuner Approach

NASA Technical Reports Server (NTRS)

Connolly, Joseph W.; Chicatelli, Amy; Garg, Sanjay

2012-01-01

This paper covers the development of a model-based engine control (MBEC) method- ology applied to an aircraft turbofan engine. Here, a linear model extracted from the Commercial Modular Aero-Propulsion System Simulation 40,000 (CMAPSS40k) at a cruise operating point serves as the engine and the on-board model. The on-board model is up- dated using an optimal tuner Kalman Filter (OTKF) estimation routine, which enables the on-board model to self-tune to account for engine performance variations. The focus here is on developing a methodology for MBEC with direct control of estimated parameters of interest such as thrust and stall margins. MBEC provides the ability for a tighter control bound of thrust over the entire life cycle of the engine that is not achievable using traditional control feedback, which uses engine pressure ratio or fan speed. CMAPSS40k is capable of modeling realistic engine performance, allowing for a verification of the MBEC tighter thrust control. In addition, investigations of using the MBEC to provide a surge limit for the controller limit logic are presented that could provide benefits over a simple acceleration schedule that is currently used in engine control architectures.

Nonlinear dynamic simulation of single- and multi-spool core engines

NASA Technical Reports Server (NTRS)

Schobeiri, T.; Lippke, C.; Abouelkheir, M.

1993-01-01

In this paper a new computational method for accurate simulation of the nonlinear dynamic behavior of single- and multi-spool core engines, turbofan engines, and power generation gas turbine engines is presented. In order to perform the simulation, a modularly structured computer code has been developed which includes individual mathematical modules representing various engine components. The generic structure of the code enables the dynamic simulation of arbitrary engine configurations ranging from single-spool thrust generation to multi-spool thrust/power generation engines under adverse dynamic operating conditions. For precise simulation of turbine and compressor components, row-by-row calculation procedures were implemented that account for the specific turbine and compressor cascade and blade geometry and characteristics. The dynamic behavior of the subject engine is calculated by solving a number of systems of partial differential equations, which describe the unsteady behavior of the individual components. In order to ensure the capability, accuracy, robustness, and reliability of the code, comprehensive critical performance assessment and validation tests were performed. As representatives, three different transient cases with single- and multi-spool thrust and power generation engines were simulated. The transient cases range from operating with a prescribed fuel schedule, to extreme load changes, to generator and turbine shut down.

Magnetized Target Fusion

NASA Technical Reports Server (NTRS)

Griffin, Steven T.

2002-01-01

Magnetized target fusion (MTF) is under consideration as a means of building a low mass, high specific impulse, and high thrust propulsion system for interplanetary travel. This unique combination is the result of the generation of a high temperature plasma by the nuclear fusion process. This plasma can then be deflected by magnetic fields to provide thrust. Fusion is initiated by a small traction of the energy generated in the magnetic coils due to the plasma's compression of the magnetic field. The power gain from a fusion reaction is such that inefficiencies due to thermal neutrons and coil losses can be overcome. Since the fusion reaction products are directly used for propulsion and the power to initiate the reaction is directly obtained from the thrust generation, no massive power supply for energy conversion is required. The result should be a low engine mass, high specific impulse and high thrust system. The key is to successfully initiate fusion as a proof-of-principle for this application. Currently MSFC is implementing MTF proof-of-principle experiments. This involves many technical details and ancillary investigations. Of these, selected pertinent issues include the properties, orientation and timing of the plasma guns and the convergence and interface development of the "pusher" plasma. Computer simulations of the target plasma's behavior under compression and the convergence and mixing of the gun plasma are under investigation. This work is to focus on the gun characterization and development as it relates to plasma initiation and repeatability.

REAR DETAIL OF RIGHT ENGINE AND WING. THRUST REVERSER REMAINS ...

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

REAR DETAIL OF RIGHT ENGINE AND WING. THRUST REVERSER REMAINS OPEN. MECHANICS JONI BAINE (R) AND BILL THEODORE(L) OPEN FLAP CARRIAGE ACCESS WITH AN IMPACT GUN. THEY WILL CHECK TRANSMISSION FLUID AND OIL THE JACK SCREW. AT FAR LEFT UTILITY MECHANICS BEGIN BODY POLISHING. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

Computational analysis of liquid hypergolic propellant rocket engines

NASA Technical Reports Server (NTRS)

Krishnan, A.; Przekwas, A. J.; Gross, K. W.

1992-01-01

The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

NASA Technical Reports Server (NTRS)

Smith, J. A.; Stechman, R. C.

1981-01-01

A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

1997-08-07

This double exposure depicts Marshall Space Flight Center's (MSFC) Test Stand 116 hosting a 60K Bantam Fastrac thrust chamber assembly test. The lower right exposure shows the engine firing in the test stand while the center exposure reveals workers monitoring the test in the interior block house of the test facility. The thrust chamber assembly is only part of the Fastrac engine project to build a low-cost engine for the X-34, an alternate light-weight unmarned launch vehicle. Both the nozzle and the engine for Fastrac are being manufactured at MSFC.

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

A complete F-1 engine assembly is shown in this photograph. Designed and developed by Rocketdye under the direction of the Marshall Space Flight Center, the engine measured 19-feet tall by 12.5 feet at the nozzle exit, and each engine produced a 1,500,000-pound thrust using liquid oxygen and kerosene as the propellant. A cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust.

Prograde evolution of the Scottish Caledonides and tectonic implications

NASA Astrophysics Data System (ADS)

Ashley, Kyle T.; Thigpen, J. Ryan; Law, Richard D.

2015-05-01

Recent thermometric analyses of samples collected in thrust-parallel structural transects across the Scandian (435-415 Ma) orogenic wedge in northwest Scotland provide a comprehensive characterization of the synorogenic retro-wedge thermal architecture. However, the paucity of petrologically-important metamorphic mineral phases (e.g., staurolite, Al-silicates) has limited investigation of pressure-temperature (P-T) histories, which hinders our ability to examine the nature of orogen-scale kinematic and thermal coupling. New data collected along a foreland-to-hinterland transect from the Moine to the Naver thrust sheets provides additional constraints for characterizing the prograde metamorphic evolution. In addition, we characterized Ti diffusion profiles in quartz inclusions in garnet to constrain duration of metamorphic heating. These results are used to develop coupled kinematic-thermal models of Scandian orogenic evolution. Early garnet core growth conditions are constrained by isopleth intersections, with peak P-T estimates determined by conventional exchange and net transfer thermobarometry and thermodynamic calculations. Most samples follow normal prograde heating and burial profiles, with peak conditions of 450 °C and 5.0 kbar in the immediate hanging wall to the Moine thrust, increasing in temperature and pressure to 733 °C and 9.5 kbar in the immediate hanging wall to the Naver thrust. These normal prograde pressure trajectories are interpreted to reflect burial of incipient thrust sheets beneath the overriding wedge at the leading edge of the orogen. Prograde heating coeval with burial is interpreted to result from surface-directed isotherm perturbation due to thrust-related advection in the overriding wedge. One exception to this is a sample from the top of the Moine thrust sheet, where prograde heating occurs during decompression (540 °C and 8.1 kbar to 590 °C and 7.0 kbar). In this case, the short lag times between motion on the Moine and Ben Hope thrusts may have limited advectionary heating until after exhumation associated with motion on the underlying Moine thrust was underway. Ti diffusion profiles in quartz inclusions in garnet suggest the near-peak thermal evolution of these rocks occurred over very short time scales (< 200,000 years). While most of the garnets are inferred to be Scandian in age, we document evidence for pre-Scandian garnet cores in structurally higher (more hinterland positioned) samples that must have grown under higher temperatures. In the hanging wall of the Moine thrust, high grossular garnets with estimated formation conditions > 9 kbar are probably of detrital origin.

Noncatalytic hydrazine thruster development - 0.050 to 5.0 pounds thrust

NASA Technical Reports Server (NTRS)

Murch, C. K.; Sackheim, R. L.; Kuenzly, J. D.; Callens, R. A.

1976-01-01

Noncatalytic (thermal-decompositon) hydrazine thrusters can operate in both the pulsing and steady-state modes to meet the propulsive requirements of long-life spacecraft. The thermal decomposition mode yields higher specific impulse than is characteristic of catalytic thrusters at similar thrust levels. This performance gain is the result of higher temperature operation and a lower fraction of ammonia dissociation. Some life limiting factors of catalytic thrusters are eliminated.

Liquid rocket booster study. Volume 2, book 4, appendices 6-8: Reports of Rocketdyne, Pratt and Whitney, and TRW

NASA Technical Reports Server (NTRS)

1988-01-01

For the pressure fed engines, detailed trade studies were conducted defining engine features such as thrust vector control methods, thrust chamber construction, etc. This was followed by engine design layouts and booster propulsion configuration layouts. For the pump fed engines parametric performance and weight data was generated for both O2/H2 and O2/RP-1 engines. Subsequent studies resulted in the selection of both LOX/RP-1 and O2/H2 propellants for the pump fed engines. More detailed analysis of the selected LOX/RP-1 and O2/H2 engines was conducted during the final phase of the study.

Structures-propulsion interactions and requirements. [large space structures

NASA Technical Reports Server (NTRS)

Coyner, J. V.

1982-01-01

The effects of low-thrust primary propulsion system characteristics on the mass, area, and orbit transfer characteristics of large space systems (LSS) were determined. Three general structural classes of LSS were considered, each with a broad range of diameters and nonstructural surface densities. While transferring the deployed structure from LEO and to GEO, an acceleration range of 0.02 to 0.1 g's was found to maximize deliverable payload based on structural mass impact. After propulsion system parametric analyses considering four propellant combinations produced values for available payload mass, length and volume, a thrust level range which maximizes deliverable LSS diameter was determined corresponding to a structure and propulsion vehicle. The engine start and/or shutdown thrust transients on the last orbit transfer (apogee) burn can impose transient loads which would be greater than the steady-state loads at the burnout acceleration. The effect of the engine thrust transients on the LSS was determined from the dynamic models upon which various engine ramps were imposed.

IR signature study of aircraft engine for variation in nozzle exit area

NASA Astrophysics Data System (ADS)

Baranwal, Nidhi; Mahulikar, Shripad P.

2016-01-01

In general, jet engines operate with choked nozzle during take-off, climb and cruise, whereas unchoking occurs while landing and taxiing (when engine is not running at full power). Appropriate thrust in an aircraft in all stages of the flight, i.e., take-off, climb, cruise, descent and landing is achieved through variation in the nozzle exit area. This paper describes the effect on thrust and IR radiance of a turbojet engine due to variation in the exit area of a just choked converging nozzle (Me = 1). The variations in the nozzle exit area result in either choking or unchoking of a just choked converging nozzle. Results for the change in nozzle exit area are analyzed in terms of thrust, mass flow rate and specific fuel consumption. The solid angle subtended (Ω) by the exhaust system is estimated analytically, for the variation in nozzle exit area (Ane), as it affects the visibility of the hot engine parts from the rear aspect. For constant design point thrust, IR radiance is studied from the boresight (ϕ = 0°, directly from the rear side) for various percentage changes in nozzle exit area (%ΔAne), in the 1.9-2.9 μm and 3-5 μm bands.

Tank 12 data dump OME integrated thrust chamber test report, phase 1

NASA Technical Reports Server (NTRS)

Pauckert, R. P.; Tobin, R. D.

1974-01-01

The test program conducted to characterize the steady state stability, thermal, and performance characteristics of the integrated thrust chamber assembly, as well as limited tests to investigate transient characteristics are described.

TRANSTRAIN: A program to compute strain transformations in composite materials

NASA Technical Reports Server (NTRS)

Ahmed, Rafiq

1990-01-01

Over the years, the solid rocket motor community has made increasing use of composite materials for thermal and structural applications. This is particularly true of solid rocket nozzles, which have used carbon phenolic and, increasingly, carbon-carbon materials to provide structural integrity and thermal protection at the high temperatures encountered during motor burn. To evaluate the degree of structural performance of nozzles and their materials and to verify analysis models, many subscale and full-scale tests are run. These provide engineers with valuable data needed to optimize design and to analyze nozzle hardware. Included among these data are strains, pressures, thrust, temperatures, and displacements. Recent nozzle test hardware has made increasing use of strain gauges embedded in the carbon composite material to measure internal strains. In order to evaluate strength, these data must be transformed into strains along the fiber directions. The fiber-direction stresses can then be calculated. A computer program written to help engineers correctly manipulate the strain data into a form that can be used to evaluate structural integrity of the nozzle is examined.

Adaptive Failure Compensation for Aircraft Flight Control Using Engine Differentials: Regulation

NASA Technical Reports Server (NTRS)

Yu, Liu; Xidong, Tang; Gang, Tao; Joshi, Suresh M.

2005-01-01

The problem of using engine thrust differentials to compensate for rudder and aileron failures in aircraft flight control is addressed in this paper in a new framework. A nonlinear aircraft model that incorporates engine di erentials in the dynamic equations is employed and linearized to describe the aircraft s longitudinal and lateral motion. In this model two engine thrusts of an aircraft can be adjusted independently so as to provide the control flexibility for rudder or aileron failure compensation. A direct adaptive compensation scheme for asymptotic regulation is developed to handle uncertain actuator failures in the linearized system. A design condition is specified to characterize the system redundancy needed for failure compensation. The adaptive regulation control scheme is applied to the linearized model of a large transport aircraft in which the longitudinal and lateral motions are coupled as the result of using engine thrust differentials. Simulation results are presented to demonstrate the effectiveness of the adaptive compensation scheme.

Performance improvements of a highly integrated digital electronic control system for an F-15 airplane

NASA Technical Reports Server (NTRS)

Putnam, T. W.; Burcham, F. W., Jr.; Andries, M. G.; Kelly, J. B.

1985-01-01

The NASA highly integrated digital electronic control (HIDEC) program is structured to conduct flight research into the benefits of integrating an aircraft flight control system with the engine control system. A brief description of the HIDEC system installed on an F-15 aircraft is provided. The adaptive engine control system (ADECS) mode is described in detail, together with simulation results and analyses that show the significant excess thrust improvements achievable with the ADECS mode. It was found that this increased thrust capability is accompanied by reduced fan stall margin and can be realized during flight conditions where engine face distortion is low. The results of analyses and simulations also show that engine thrust response is improved and that fuel consumption can be reduced. Although the performance benefits that accrue because of airframe and engine control integration are being demonstrated on an F-15 aircraft, the principles are applicable to advanced aircraft such as the advanced tactical fighter and advanced tactical aircraft.

EC94-42645-9

NASA Image and Video Library

1994-06-27

The modified F-18 High Alpha Research Vehicle (HARV) carries out air flow studies on a flight from the Dryden Flight Research Center, Edwards, California. Using oil, researchers were able to track the air flow across the wing at different speeds and angles of attack. A thrust vectoring system had been installed on the engines' exhaust nozzles for the high angle of attack research program. The thrust vectoring system, linked to the aircraft's flight control system, moves a set of three paddles on each engine to redirect thrust for directional control and increased maneuverability at angles of attack at up to 70 degrees.

Magnetohydrodynamic Augmented Propulsion Experiment

NASA Technical Reports Server (NTRS)

Litchford, Ron J.

2008-01-01

Over the past several years, efforts have been under way to design and develop an operationally flexible research facility for investigating the use of cross-field MHD accelerators as a potential thrust augmentation device for thermal propulsion systems. The baseline configuration for this high-power experimental facility utilizes a 1.5-MWe multi-gas arc-heater as a thermal driver for a 2-MWe MHD accelerator, which resides in a large-bore 2-tesla electromagnet. A preliminary design study using NaK seeded nitrogen as the working fluid led to an externally diagonalized segmented MHD channel configuration based on an expendable heat-sink design concept. The current status report includes a review of engineering/design work and performance optimization analyses and summarizes component hardware fabrication and development efforts, preliminary testing results, and recent progress toward full-up assembly and testing

Status of Magnetohydrodynamic Augmented Propulsion Experiment

NASA Technical Reports Server (NTRS)

Litchford, Ron J.; Lineberry, John T.

2007-01-01

Over the past several years, efforts have been under way to design and develop an operationally flexible research facility for investigating the use of cross-field MHD accelerators as a potential thrust augmentation device for thermal propulsion systems, The baseline configuration for this high-power experimental facility utilizes a 1,5-MW, multi-gas arc-heater as a thermal driver for a 2-MW, MHD accelerator, which resides in a large-bore 2-tesla electromagnet. A preliminary design study using NaK seeded nitrogen as the working fluid led to an externally diagonalized segmented MHD channel configuration based on an expendable beat-sink design concept. The current status report includes a review of engineering/design work and performance optimization analyses and summarizes component hardware fabrication and development efforts, preliminary testing results, and recent progress toward full-up assembly and testing

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.

High-power, null-type, inverted pendulum thrust stand.

PubMed

Xu, Kunning G; Walker, Mitchell L R

2009-05-01

This article presents the theory and operation of a null-type, inverted pendulum thrust stand. The thrust stand design supports thrusters having a total mass up to 250 kg and measures thrust over a range of 1 mN to 5 N. The design uses a conventional inverted pendulum to increase sensitivity, coupled with a null-type feature to eliminate thrust alignment error due to deflection of thrust. The thrust stand position serves as the input to the null-circuit feedback control system and the output is the current to an electromagnetic actuator. Mechanical oscillations are actively damped with an electromagnetic damper. A closed-loop inclination system levels the stand while an active cooling system minimizes thermal effects. The thrust stand incorporates an in situ calibration rig. The thrust of a 3.4 kW Hall thruster is measured for thrust levels up to 230 mN. The uncertainty of the thrust measurements in this experiment is +/-0.6%, determined by examination of the hysteresis, drift of the zero offset and calibration slope variation.

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hulka, James; Yang, Virog

2009-01-01

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Emergency Multiengine Aircraft System for Lateral Control Using Differential Thrust Control of Wing Engines

NASA Technical Reports Server (NTRS)

Burken, John J. (Inventor); Burcham, Frank W., Jr. (Inventor); Bull, John (Inventor)

2000-01-01

Development of an emergency flight control system is disclosed for lateral control using only differential engine thrust modulation of multiengine aircraft is currently underway. The multiengine has at least two engines laterally displaced to the left and right from the axis of the aircraft. In response to a heading angle command psi(sub c) is to be tracked. By continually sensing the heading angle psi of the aircraft and computing a heading error signal psi(sub e) as a function of the difference between the heading angle command psi(sub c) and the sensed heading angle psi, a track control signal is developed with compensation as a function of sensed bank angle phi. Bank angle rate phi, or roll rate p, yaw rate tau, and true velocity produce an aircraft thrust control signal ATC(sub psi(L,R)). The thrust control signal is differentially applied to the left and right engines, with equal amplitude and opposite sign, such that a negative sign is applied to the control signal on the side of the aircraft. A turn is required to reduce the error signal until the heading feedback reduces the error to zero.

A graphite-lined regeneratively cooled thrust chamber

NASA Technical Reports Server (NTRS)

Stubbs, V. R.

1972-01-01

Design concepts, based on use of graphite as a thermal barrier for regeneratively cooled FLOX-methane thrust chambers, have been screened and concepts selected for detailed thermodynamic, stress, and fabrication analyses. A single design employing AGCarb-101, a fibrous graphite composite material, for a thermal barrier liner and an electroformed nickel structure with integral coolant passages was selected for fabrication and testing. The fabrication processes and the test results are described and illustrated.

The 30-centimeter ion thrust subsystem design manual

NASA Technical Reports Server (NTRS)

1979-01-01

The principal characteristics of the 30-centimeter ion propulsion thrust subsystem technology that was developed to satisfy the propulsion needs of future planetary and early orbital missions are described. Functional requirements and descriptions, interface and performance requirements, and physical characteristics of the hardware are described at the thrust subsystem, BIMOD engine system, and component level.

Spectroscopy-based thrust sensor for high-speed gaseous flows

NASA Technical Reports Server (NTRS)

Hanson, Ronald K. (Inventor)

1993-01-01

A system and method for non-intrusively obtaining the thrust value of combustion by-products of a jet engine is disclosed herein. The system includes laser elements for inducing absorption for use in determining the axial velocity and density of the jet flow stream and elements for calculating the thrust value therefrom.

An experimental investigation of thrust vectoring two-dimensional convergent-divergent nozzles installed in a twin-engine fighter model at high angles of attack

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Mason, Mary L.; Leavitt, Laurence D.

1990-01-01

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine thrust vectoring capability of subscale 2-D convergent-divergent exhaust nozzles installed on a twin engine general research fighter model. Pitch thrust vectoring was accomplished by downward rotation of nozzle upper and lower flaps. The effects of nozzle sidewall cutback were studied for both unvectored and pitch vectored nozzles. A single cutback sidewall was employed for yaw thrust vectoring. This investigation was conducted at Mach numbers ranging from 0 to 1.20 and at angles of attack from -2 to 35 deg. High pressure air was used to simulate jet exhaust and provide values of nozzle pressure ratio up to 9.

Over-the-wing model thrust reverser noise tests

NASA Technical Reports Server (NTRS)

Goodykoontz, J.; Gutierrez, O.

1977-01-01

Static acoustic tests were conducted on a 1/12 scale model over-the-wing target type thrust reverser. The model configuration simulates a design that is applicable to the over-the-wing short-haul advanced technology engine. Aerodynamic screening tests of a variety of reverser designs identified configurations that satisfied a reverse thrust requirement of 35 percent of forward thrust at a nozzle pressure ratio of 1.29. The variations in the reverser configuration included, blocker door angle, blocker door lip angle and shape, and side skirt shape. Acoustic data are presented and compared for the various configurations. The model data scaled to a single full size engine show that peak free field perceived noise (PN) levels at a 152.4 meter sideline distance range from 98 to 104 PNdb.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accomodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part I-Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2-Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25.800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3-Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine, two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5,000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC material in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accommodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part 1: Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2: Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25,800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3: Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine. two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5.000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC materials in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Micronozzles: 3D numerical structural and gas dynamics modeling, fabrication, and preliminary experimental results

NASA Astrophysics Data System (ADS)

Borovkov, Alexei I.; Pyatishev, Evgenij N.; Lurie, Mihail S.; Korshunov, Andrey V.; Akulshin, Y. D.; Dolganov, A. G.; Sabadash, V. O.

2000-02-01

The tiny engines, founded on the principle of reactive thrust, are one of most perspective actuators developed by modern micromechanics. These engines can be applied for such apparent problems, as orientation and stabilization of small space objects, but also as local or distributed reactive thrust of new phylum of aerospace objects, for control of boundary layer of flying objects and in series of converting power devices of different purposes. Distinctive features of jet tiny engines are profitability (very large thrust-to-weight ratio) and high (milliseconds) response, which makes them to irreplaceable elements in control systems and, specially, in distributed power generations. These features are provided the minimum sizes, high pressure in working chambers and hypersonic velocity of propulsive jet. Topologically micronozzles are designed as the flat batch devices (3 layers as minimum). The lower and upper layers make flat walls of the nozzle and mainly influence on strength properties of the device. The mean layer reshapes geometry and determines gas dynamic characteristic of the nozzle. A special problem is the opening-up of the combustion-mixture, which is not esteemed in this work. It is necessary to allow for effect of considerable local stresses arising at the expense of static and dynamic loading at design of the jet tiny engines. Thermal gas dynamic processes in the chamber and nozzle determine the values and nature of these stresses, which are hardly studied for the microdevices. The priority is mathematical and experimental simulation of these processes. The most suitable object for initial phase of experimental simulation is the 'cold' engine. The demanded chamber static pressure is formed by external compressed air. In Laboratory of Microtechnology and MicroElectroMechanical Systems a number of such tiny engines with different shapes of the chamber's and the nozzles' surfaces were designed, made and tested. The engines were produced from photosensing glass by methods of microtechnology on the basis of photolithography processes. After expositing through a mask the latent map of the glass was 'showed' by heat treatment and etched. The obtained parts sitallized and subjected to level-by-level assembly. At experiments on 'ardent' engines it is supposed to keep the basic stages of a technological route, but to use stronger and temperature- resistant materials including coating from high-strength membranes plotted by vacuum deposition methods. During trial tests, for the 'cold' engine with an altitude of a nozzle of 1.2 mm and width of the throat of 0.4 mm at chamber pressure 0.6 MPa the exhaust velocity on escaping of the nozzle about 1.5 M was obtained. The engine thrust has compounded 45 gr. The obtained data are in satisfactory conformity with 1D computation and allow to proceed piloting objects of other range of the characteristics. The microactuators having high response and profitability are demanded for perspective small aerospace objects. This activators are indispensable for creation of distributed thrust and control of boundary layer of micro air flying objects (MAV), for devices of stabilization and orientation of micro-satellites. A number of such activators forms on the areas of flat micronozzle devices. Developed micronozzles should provide demanded parameters at the expense of a high level of pressure in working chamber and supersonic exhaust velocities. At creation of the micronozzle the effect of considerable loads arising as at the expense of static, and dynamic loading should be mentioned. Thermomechanics-gasodynamic processes in the chamber and nozzle determine the nature and kind of loading. Mathematical and experimental simulation of these hardly studied for the microscopic object processes is necessary.

Micronozzles: 3D numerical structural and gas dynamics modeling, fabrication, and preliminary experimental results

NASA Astrophysics Data System (ADS)

Borovkov, Alexei I.; Pyatishev, Evgenij N.; Lurie, Mihail S.; Korshunov, Andrey V.; Akulshin, Y. D.; Dolganov, A. G.; Sabadash, V. O.

2001-02-01

The tiny engines, founded on the principle of reactive thrust, are one of most perspective actuators developed by modern micromechanics. These engines can be applied for such apparent problems, as orientation and stabilization of small space objects, but also as local or distributed reactive thrust of new phylum of aerospace objects, for control of boundary layer of flying objects and in series of converting power devices of different purposes. Distinctive features of jet tiny engines are profitability (very large thrust-to-weight ratio) and high (milliseconds) response, which makes them to irreplaceable elements in control systems and, specially, in distributed power generations. These features are provided the minimum sizes, high pressure in working chambers and hypersonic velocity of propulsive jet. Topologically micronozzles are designed as the flat batch devices (3 layers as minimum). The lower and upper layers make flat walls of the nozzle and mainly influence on strength properties of the device. The mean layer reshapes geometry and determines gas dynamic characteristic of the nozzle. A special problem is the opening-up of the combustion-mixture, which is not esteemed in this work. It is necessary to allow for effect of considerable local stresses arising at the expense of static and dynamic loading at design of the jet tiny engines. Thermal gas dynamic processes in the chamber and nozzle determine the values and nature of these stresses, which are hardly studied for the microdevices. The priority is mathematical and experimental simulation of these processes. The most suitable object for initial phase of experimental simulation is the 'cold' engine. The demanded chamber static pressure is formed by external compressed air. In Laboratory of Microtechnology and MicroElectroMechanical Systems a number of such tiny engines with different shapes of the chamber's and the nozzles' surfaces were designed, made and tested. The engines were produced from photosensing glass by methods of microtechnology on the basis of photolithography processes. After expositing through a mask the latent map of the glass was 'showed' by heat treatment and etched. The obtained parts sitallized and subjected to level-by-level assembly. At experiments on 'ardent' engines it is supposed to keep the basic stages of a technological route, but to use stronger and temperature- resistant materials including coating from high-strength membranes plotted by vacuum deposition methods. During trial tests, for the 'cold' engine with an altitude of a nozzle of 1.2 mm and width of the throat of 0.4 mm at chamber pressure 0.6 MPa the exhaust velocity on escaping of the nozzle about 1.5 M was obtained. The engine thrust has compounded 45 gr. The obtained data are in satisfactory conformity with 1D computation and allow to proceed piloting objects of other range of the characteristics. The microactuators having high response and profitability are demanded for perspective small aerospace objects. This activators are indispensable for creation of distributed thrust and control of boundary layer of micro air flying objects (MAV), for devices of stabilization and orientation of micro-satellites. A number of such activators forms on the areas of flat micronozzle devices. Developed micronozzles should provide demanded parameters at the expense of a high level of pressure in working chamber and supersonic exhaust velocities. At creation of the micronozzle the effect of considerable loads arising as at the expense of static, and dynamic loading should be mentioned. Thermomechanics-gasodynamic processes in the chamber and nozzle determine the nature and kind of loading. Mathematical and experimental simulation of these hardly studied for the microscopic object processes is necessary.

Performance seeking control (PSC) for the F-15 highly integrated digital electronic control (HIDEC) aircraft

NASA Technical Reports Server (NTRS)

Orme, John S.

1995-01-01

The performance seeking control algorithm optimizes total propulsion system performance. This adaptive, model-based optimization algorithm has been successfully flight demonstrated on two engines with differing levels of degradation. Models of the engine, nozzle, and inlet produce reliable, accurate estimates of engine performance. But, because of an observability problem, component levels of degradation cannot be accurately determined. Depending on engine-specific operating characteristics PSC achieves various levels performance improvement. For example, engines with more deterioration typically operate at higher turbine temperatures than less deteriorated engines. Thus when the PSC maximum thrust mode is applied, for example, there will be less temperature margin available to be traded for increasing thrust.

Low Thrust Orbital Maneuvers Using Ion Propulsion

NASA Astrophysics Data System (ADS)

Ramesh, Eric

2011-10-01

Low-thrust maneuver options, such as electric propulsion, offer specific challenges within mission-level Modeling, Simulation, and Analysis (MS&A) tools. This project seeks to transition techniques for simulating low-thrust maneuvers from detailed engineering level simulations such as AGI's Satellite ToolKit (STK) Astrogator to mission level simulations such as the System Effectiveness Analysis Simulation (SEAS). Our project goals are as follows: A) Assess different low-thrust options to achieve various orbital changes; B) Compare such approaches to more conventional, high-thrust profiles; C) Compare computational cost and accuracy of various approaches to calculate and simulate low-thrust maneuvers; D) Recommend methods for implementing low-thrust maneuvers in high-level mission simulations; E) prototype recommended solutions.

Human Exploration and Settlement of the Moon Using LUNOX-Augmented NTR Propulsion

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Culver, Donald W.; Bulman, Melvin J.

1995-01-01

An innovative trimodal nuclear thermal rocket (NTR) concept is described which combines conventional liquid hydrogen (LH2)-cooled NTR, Brayton cycle power generation and supersonic combustion ramjet (scramjet) technologies. Known as the liquid oxygen (LOX) augmented NTR (LANTR), this concept utilizes the large divergent section of the NTR nozzle as an 'afterburner' into which LOX is injected and supersonically combusted with nuclear preheated hydrogen emerging from the LANTR's choked sonic throat--'scramjet propulsion in reverse.' By varying the oxygen-to-hydrogen mixture ratio (MR), the LANTR can operate over a wide range of thrust and specific impulse (Isp) values while the reactor core power level remains relatively constant. As the MR varies from zero to seven, the thrust-to-weight ratio for a 15 thousand pound force (klbf) NTR increases by approximately 440%--from 3 to 13--while the Isp decreases by only approximately 45%--from 940 to 515 seconds. This thrust augmentation feature of the LANTR means that 'big engine' performance can be obtained using smaller more affordable, easier to test NTR engines. 'Reoxidizing' the bipropellant LANTR system in low lunar orbit (LLO) with high density 'lunar-derived' LOX (LUNOX) enables a reusable, reduced size and mass lunar transfer vehicle (LTV) which can be deployed and resupplied using two 66 t-class Shuttle-derived launch vehicles. The reusable LANTR can also transport 200 to 300% more payload on each piloted round trip mission than an expendable 'all LH2' NTR system. As initial outposts grow to eventual lunar settlements and LUNOX production capacity increases, the LANTR concept can also enable a rapid 'commuter' shuttle capable of 36 to 24 hour 'one way' trips to the Moon and back with reasonable size vehicles and initial mass in low Earth orbit (IMLEO) requirements.

Human exploration and settlement of the Moon using LUNOX-augmented NTR propulsion

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.; Culver, Donald W.; Bulman, Melvin J.

1995-10-01

An innovative trimodal nuclear thermal rocket (NTR) concept is described which combines conventional liquid hydrogen (LH2)-cooled NTR, Brayton cycle power generation and supersonic combustion ramjet (scramjet) technologies. Known as the liquid oxygen (LOX) augmented NTR (LANTR), this concept utilizes the large divergent section of the NTR nozzle as an 'afterburner' into which LOX is injected and supersonically combusted with nuclear preheated hydrogen emerging from the LANTR's choked sonic throat--'scramjet propulsion in reverse.' By varying the oxygen-to-hydrogen mixture ratio (MR), the LANTR can operate over a wide range of thrust and specific impulse (Isp) values while the reactor core power level remains relatively constant. As the MR varies from zero to seven, the thrust-to-weight ratio for a 15 thousand pound force (klbf) NTR increases by approximately 440%--from 3 to 13--while the Isp decreases by only approximately 45%--from 940 to 515 seconds. This thrust augmentation feature of the LANTR means that 'big engine' performance can be obtained using smaller more affordable, easier to test NTR engines. 'Reoxidizing' the bipropellant LANTR system in low lunar orbit (LLO) with high density 'lunar-derived' LOX (LUNOX) enables a reusable, reduced size and mass lunar transfer vehicle (LTV) which can be deployed and resupplied using two 66 t-class Shuttle-derived launch vehicles. The reusable LANTR can also transport 200 to 300% more payload on each piloted round trip mission than an expendable 'all LH2' NTR system. As initial outposts grow to eventual lunar settlements and LUNOX production capacity increases, the LANTR concept can also enable a rapid 'commuter' shuttle capable of 36 to 24 hour 'one way' trips to the Moon and back with reasonable size vehicles and initial mass in low Earth orbit (IMLEO) requirements.

Human exploration and settlement of the moon using lunox-augmented NTR propulsion

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.; Culver, Donald W.; Bulman, Melvin J.

1995-01-01

An innovative trimodal nuclear thermal rocket (NTR) concept is described which combines conventional liquid hydrogen (LH2)-cooled NTR, Brayton cycle power generation and supersonic combustion ramjet (scramjet) technologies. Known as the liquid oxygen (LOS)-augmented NTR (LANTR), this concept utilizes the large divergent section of the NTR nozzle as an ``afterburner'' into which LOX is injected and supersonically combusted with nuclear preheated hydrogen emerging from the LANTR's choked sonic throat—``scramjet propulsion in reverse.'' By varying the oxygen-to-hydrogen mixture ratio (MR), the LANTR can operate over a wide range of thrust and specific impulse (Isp) values while the reactor core power level remains relatively constant. As the MR varies from zero to seven, the thrust-to-weight ratio for a 15 thousand pound force (klbf) NTR increases by ˜440%—from 3 to 13—while the Isp decreases by only ˜45%—from 940 to 515 seconds. This thrust augmentation feature of the LANTR means that ``big engine'' performance can be obtained using smaller, more affordable, easier to test NTR engines. ``Reoxidizing'' the bipropellant LANTR system in low lunar orbit (LLO) with high density ``lunar-derived'' LOX (LUNOX) enables a reusable, reduced size and mass lunar transfer vehicle (LTV) which can be deployed and resupplied using two 66 t-class Shuttle-derived launch vehicles. The reusable LANTR can also transport 200 to 300% more payload on each piloted round trip mission than an expendable ``all LH2'' NTR system. As initial outposts grow to eventual lunar settlements and LUNOX production capacity increases, the LANTR concept can also enable a rapid ``commuter'' shuttle capable of 36 to 24 hour ``one way'' trip to the Moon and back with reasonable size vehicles and initial mass in low Earth orbit (IMLEO) requirements.

Implementation of the Orbital Maneuvering Systems Engine and Thrust Vector Control for the European Service Module

NASA Technical Reports Server (NTRS)

Millard, Jon

2014-01-01

The European Space Agency (ESA) has entered into a partnership with the National Aeronautics and Space Administration (NASA) to develop and provide the Service Module (SM) for the Orion Multipurpose Crew Vehicle (MPCV) Program. The European Service Module (ESM) will provide main engine thrust by utilizing the Space Shuttle Program Orbital Maneuvering System Engine (OMS-E). Thrust Vector Control (TVC) of the OMS-E will be provided by the Orbital Maneuvering System (OMS) TVC, also used during the Space Shuttle Program. NASA will be providing the OMS-E and OMS TVC to ESA as Government Furnished Equipment (GFE) to integrate into the ESM. This presentation will describe the OMS-E and OMS TVC and discuss the implementation of the hardware for the ESM.

Evaluation of a simplified gross thrust calculation technique using two prototype F100 turbofan engines in an altitude facility

NASA Technical Reports Server (NTRS)

Kurtenbach, F. J.

1979-01-01

The technique which relies on afterburner duct pressure measurements and empirical corrections to an ideal one dimensional flow analysis to determine thrust is presented. A comparison of the calculated and facility measured thrust values is reported. The simplified model with the engine manufacturer's gas generator model are compared. The evaluation was conducted over a range of Mach numbers from 0.80 to 2.00 and at altitudes from 4020 meters to 15,240 meters. The effects of variations in inlet total temperature from standard day conditions were explored. Engine conditions were varied from those normally scheduled for flight. The technique was found to be accurate to a twice standard deviation of 2.89 percent, with accuracy a strong function of afterburner duct pressure difference.

The effects of compressor seventh-stage bleed air extraction on performance of the F100-PW-220 afterburning turbofan engine

NASA Technical Reports Server (NTRS)

Evans, Alison B.

1991-01-01

A study was conducted to determine the effects of seventh-stage compressor bleed on the performance of the F100 afterburning turbofan engine. The effects of bleed on thrust, specific fuel consumption, fan turbine inlet temperature, bleed total pressure, and bleed total temperature were obtained from the engine manufacturer's status deck computer simulation. These effects were determined for power settings of intermediate, partial afterburning, and maximum afterburning for Mach numbers between 0.6 and 2.2 and for altitudes of 30,000, 40,000, and 50,000 ft. It was found that thrust loss and specific fuel consumption increase were approximately linear functions of bleed flow and, based on a percent-thrust change basis, were approximately independent of power setting.

Interrogation of possible imaging conditions for radiation sensitive metal organic frameworks in transmission electron microscopes

NASA Astrophysics Data System (ADS)

Patel, Harinkumar Rajendrabhai

One of the main area of research currently in air-breathing propulsion is increasing the fuel efficiency of engines. Increasing fuel efficiency of an air-breathing engine will be advantageous for civil transport as well as military aircraft. This objective can be achieved in several ways. Present design models are developed based on their uses: commercial transport, high range rescue aircraft, military aircraft. One of the main property of military aircraft is possessing high thrust but increasing fuel efficiency will also be advantageous resulting in more time in combat. Today's engine design operates best at their design point and has reduced thrust and high fuel consumption values in off-design. The adaptive cycle engine concept was introduced to overcome this problem. The adaptive cycle engine is a variable cycle engine concept equipped with an extra bypass (3rd bypass) stream. This engine varies the bypass ratio and the fan pressure ratio, the two main parameters affecting thrust and fuel consumption values of the engine. In cruise, more flow will flow through the third stream resulting in the high bypass engine giving lower fuel consumption. on the other hand, the engine will act as a low bypass engine producing more thrust by allowing more air to flow through core while in combat. The simulation of this engine was carried out using the Numerical Propulsion System Simulation (NPSS) software. The effect of the bypass ratio and the fan pressure ratio along with Mach number were studied. After the parametric variation study, the mixture configuration was also studied. Once the effect of the parameters were understood, the best design operating point configuration was selected and then the engine performance for off-design was calculated. Optimum values of bypass ratio and fan pressure ratio were also obtained for each altitude selected for off-design performance.

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.

Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems

NASA Astrophysics Data System (ADS)

Ryzhkov, V.; Lapshin, E.

2018-01-01

The paper presents an algorithm for designing liquid small thrust rocket engine, the process of which consists of five aggregated stages with feedback. Three stages of the algorithm provide engineering support for design, and two stages - the actual engine design. A distinctive feature of the proposed approach is a deep study of the main technical solutions at the stage of engineering analysis and interaction with the created knowledge (data) base, which accelerates the process and provides enhanced design quality. The using multifunctional graphic package Siemens NX allows to obtain the final product -rocket engine and a set of design documentation in a fairly short time; the engine design does not require a long experimental development.

Design and Demonstration of Emergency Control Modes for Enhanced Engine Performance

NASA Technical Reports Server (NTRS)

Liu, Yuan; Litt, Jonathan S.; Guo, Ten-Huei

2013-01-01

A design concept is presented for developing control modes that enhance aircraft engine performance during emergency flight scenarios. The benefits of increased engine performance to overall vehicle survivability during these situations may outweigh the accompanied elevated risk of engine failure. The objective involves building control logic that can consistently increase engine performance beyond designed maximum levels based on an allowable heightened probability of failure. This concept is applied to two previously developed control modes: an overthrust mode that increases maximum engine thrust output and a faster response mode that improves thrust response to dynamic throttle commands. This paper describes the redesign of these control modes and presents simulation results demonstrating both enhanced engine performance and robust maintenance of the desired elevated risk level.

Test and evaluation of the HIDEC engine uptrim algorithm

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemented into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

Reaction Control Engine for Space Launch Initiative

NASA Technical Reports Server (NTRS)

2002-01-01

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Model Test of the Aerospace Laser Propulsion Engine

NASA Astrophysics Data System (ADS)

Ageichik, Alexander A.; Egorov, Maxim S.; Ostapenko, Svetlana V.; Rezunkov, Yuri A.; Safronov, Alexander L.; Stepanov, Vladimir V.

2005-04-01

One of the main results of the experimental and theoretical investigations made under the ISTC Project ♯ 1801 is the original design of Aerospace Laser Propulsion Engine (ASLPE) developed. The designed characteristics of the ASLPE flight model are experimentally approved, including the test experiments with a solid propellant. The obtained momentum coupling coefficient is rather high and comparable one with respect to the coefficient obtained by other researchers. Moreover, it is experimentally demonstrated that the thrust characteristics of the ASLPE flight model does not depend on angular aberrations of the beam coming onto the beam concentrator of the model with the incident angle of 0.01 radian. The experiments also demonstrated that successful launching of the vehicle with the ASLPE under the laser characteristics is possible also if the vehicle mass will be decreased and the thermal blooming effect will be eliminated.

Space shuttle orbit maneuvering engine, reusable thrust chamber program. Task 6: Data dump hot fuel element investigation

NASA Technical Reports Server (NTRS)

Nurick, W. H.

1974-01-01

An evaluation of reusable thrust chambers for the space shuttle orbit maneuvering engine was conducted. Tests were conducted using subscale injector hot-fire procedures for the injector configurations designed for a regenerative cooled engine. The effect of operating conditions and fuel temperature on combustion chamber performance was determined. Specific objectives of the evaluation were to examine the optimum like-doublet element geometry for operation at conditions consistent with a fuel regeneratively cooled engine (hot fuel, 200 to 250 F) and the sensitivity of the triplet injector element to hot fuels.

The pasty propellant rocket engine development

NASA Astrophysics Data System (ADS)

Kukushkin, V. I.; Ivanchenko, A. N.

1993-06-01

The paper describes a newly developed pasty propellant rocket engine (PPRE) and the combustion process and presents results of performance tests. It is shown that, compared with liquid propellant rocket engines, the PPREs can regulate the thrust level within a wider range, are safer ecologically, and have better weight characteristics. Compared with solid propellant rocket engines, the PPREs may be produced with lower costs and more safely, are able to regulate thrust performance within a wider range, and are able to offer a greater scope for the variation of the formulation components and propellant characteristics. Diagrams of the PPRE are included.

Advanced engine study program

NASA Technical Reports Server (NTRS)

Masters, A. I.; Galler, D. E.; Denman, T. F.; Shied, R. A.; Black, J. R.; Fierstein, A. R.; Clark, G. L.; Branstrom, B. R.

1993-01-01

A design and analysis study was conducted to provide advanced engine descriptions and parametric data for space transfer vehicles. The study was based on an advanced oxygen/hydrogen engine in the 7,500 to 50,000 lbf thrust range. Emphasis was placed on defining requirements for high-performance engines capable of achieving reliable and versatile operation in a space environment. Four variations on the expander cycle were compared, and the advantages and disadvantages of each were assessed. Parametric weight, envelope, and performance data were generated over a range of 7,500 to 50,000 lb thrust and a wide range of chamber pressure and nozzle expansion ratio.

Space shuttle orbit maneuvering engine reusable thrust chamber

NASA Technical Reports Server (NTRS)

1972-01-01

A data dump is presented containing space shuttle orbiter maneuvering engine performance, weight, envelope, and interface pressure requirements for candidate propellant combinations (NTO/MMH, NTO50-50, LOX/MMH, LOX/50-50, LOX/N2H4, LOX/C3H8, and LOX/RP-1) and cooling concepts (regenerative and dump/film). These data are presented parametrically for the thrust, chamber pressure, nozzle expansion ratio, and engine mixture ratio ranges of interest. Also included is information describing sensitivity to system changes; reliability, maintainability and safety; development programs and associated critical technology areas; engine cost comparisons during development and operation; and ecological effects.

Research Technology

NASA Image and Video Library

2002-03-13

NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama, has begun a series of engine tests on the Reaction Control Engine developed by TRW Space and Electronics for NASA's Space Launch Initiative (SLI). SLI is a technology development effort aimed at improving the safety, reliability, and cost effectiveness of space travel for reusable launch vehicles. The engine in this photo, the first engine tested at MSFC that includes SLI technology, was tested for two seconds at a chamber pressure of 185 pounds per square inch absolute (psia). Propellants used were liquid oxygen as an oxidizer and liquid hydrogen as fuel. Designed to maneuver vehicles in orbit, the engine is used as an auxiliary propulsion system for docking, reentry, fine-pointing, and orbit transfer while the vehicle is in orbit. The Reaction Control Engine has two unique features. It uses nontoxic chemicals as propellants, which creates a safer environment with less maintenance and quicker turnaround time between missions, and it operates in dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The force of low level thrust allows the vehicle to fine-point maneuver and dock, while the force of the high level thrust is used for reentry, orbital transfer, and course positioning.

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.

High-Performance Bipropellant Engine

NASA Technical Reports Server (NTRS)

Biaglow, James A.; Schneider, Steven J.

1999-01-01

TRW, under contract to the NASA Lewis Research Center, has successfully completed over 10 000 sec of testing of a rhenium thrust chamber manufactured via a new-generation powder metallurgy. High performance was achieved for two different propellants, N2O4- N2H4 and N2O4 -MMH. TRW conducted 44 tests with N2O4-N2H4, accumulating 5230 sec of operating time with maximum burn times of 600 sec and a specific impulse Isp of 333 sec. Seventeen tests were conducted with N2O4-MMH for an additional 4789 sec and a maximum Isp of 324 sec, with a maximum firing duration of 700 sec. Together, the 61 tests totalled 10 019 sec of operating time, with the chamber remaining in excellent condition. Of these tests, 11 lasted 600 to 700 sec. The performance of radiation-cooled rocket engines is limited by their operating temperature. For the past two to three decades, the majority of radiation-cooled rockets were composed of a high-temperature niobium alloy (C103) with a disilicide oxide coating (R512) for oxidation resistance. The R512 coating practically limits the operating temperature to 1370 C. For the Earth-storable bipropellants commonly used in satellite and spacecraft propulsion systems, a significant amount of fuel film cooling is needed. The large film-cooling requirement extracts a large penalty in performance from incomplete mixing and combustion. A material system with a higher temperature capability has been matured to the point where engines are being readied for flight, particularly the 100-lb-thrust class engine. This system has powder rhenium (Re) as a substrate material with an iridium (Ir) oxidation-resistant coating. Again, the operating temperature is limited by the coating; however, Ir is capable of long-life operation at 2200 C. For Earth-storable bipropellants, this allows for the virtual elimination of fuel film cooling (some film cooling is used for thermal control of the head end). This has resulted in significant increases in specific impulse performance (15 to 20 sec). To determine the merits of a powder rhenium thrust chamber, Lewis On-Board Propulsion Branch directed TRW (under the Space Storable Rocket Technology Program and the High Pressure Earth Storable Rocket Technology Program) to design, fabricate, and test an engineering model to serve as a technology demonstrator.

NTREES Testing and Operations Status

NASA Technical Reports Server (NTRS)

Emrich, Bill

2007-01-01

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today's best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. Currently, the construction of such a simulator has been completed at the Marshall Space Flight Center, and will be used in the future to evaluate a wide variety of fuel element designs and the materials of which they are fabricated. This present work addresses the operational status of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Investigation on Composite Throat Insert For Cryogenic Engines

NASA Astrophysics Data System (ADS)

Ayyappan, G.; Tiwari, S. B.; Praveen, RS; Mohankumar, L.; Jathaveda, M.; Ganesh, P.

2017-02-01

Injector element testing is an important step in the development and qualification of the cryogenic rocket engines. For the purpose of characterising the injectors, sub scale chambers are used. In order to assess the performance of the injectors, different configurations of the injectors are tested using a combustion chamber and a convergent-divergent nozzle. Pressure distribution along the wall of the chamber and throat insert is obtained from the CFD analysis and temperature distribution is obtained from thermal analysis. Thermo-structural analysis is carried out for the sub-scale model of throat inert using temperature dependent material properties. For the experiments a sub-scale model of the thrust chamber is realised. Injector element tests are carried out for the studies. The objective of the present study is to investigate the behaviour of different throat inserts, mainly graphite, 2-D Carbon-Carbon(2D C-C), 4-D Carbon-Carbon (4D C-C) and Silica Phenolic (SP), under pressure and thermal load for repeated operation of the engine. Analytical results are compared with the test results. The paper gives the results of theoretical studies and experiments conducted with all the four type of throat material. It is concluded that 2D C-C is superior in terms of throat erosion being the least under specified combustion environment.

Design and evaluation of thrust vectored nozzles using a multicomponent thrust stand

NASA Technical Reports Server (NTRS)

Carpenter, Thomas W.; Blattner, Ernest W.; Stagner, Robert E.; Contreras, Juanita; Lencioni, Dennis; Mcintosh, Greg

1990-01-01

Future aircraft with the capability of short takeoff and landing, and improved maneuverability especially in the post-stall flight regime will incorporate exhaust nozzles which can be thrust vectored. In order to conduct thrust vector research in the Mechanical Engineering Department at Cal Poly, a program was planned with two objectives; design and construct a multicomponent thrust stand for the specific purpose of measuring nozzle thrust vectors; and to provide quality low moisture air to the thrust stand for cold flow nozzle tests. The design and fabrication of the six-component thrust stand was completed. Detailed evaluation tests of the thrust stand will continue upon the receipt of one signal conditioning option (-702) for the Fluke Data Acquisition System. Preliminary design of thrust nozzles with air supply plenums were completed. The air supply was analyzed with regard to head loss. Initial flow visualization tests were conducted using dual water jets.

Development of a two-dimensional dual pendulum thrust stand for Hall thrusters.

PubMed

Nagao, N; Yokota, S; Komurasaki, K; Arakawa, Y

2007-11-01

A two-dimensional dual pendulum thrust stand was developed to measure thrust vectors [axial and horizontal (transverse) direction thrusts] of a Hall thruster. A thruster with a steering mechanism is mounted on the inner pendulum, and thrust is measured from the displacement between inner and outer pendulums, by which a thermal drift effect is canceled out. Two crossover knife-edges support each pendulum arm: one is set on the other at a right angle. They enable the pendulums to swing in two directions. Thrust calibration using a pulley and weight system showed that the measurement errors were less than 0.25 mN (1.4%) in the main thrust direction and 0.09 mN (1.4%) in its transverse direction. The thrust angle of the thrust vector was measured with the stand using the thruster. Consequently, a vector deviation from the main thrust direction of +/-2.3 degrees was measured with the error of +/-0.2 degrees under the typical operating conditions for the thruster.

Metaheuristic and Machine Learning Models for TFE-731-2, PW4056, and JT8D-9 Cruise Thrust

NASA Astrophysics Data System (ADS)

Baklacioglu, Tolga

2017-08-01

The requirement for an accurate engine thrust model has a major antecedence in airline fuel saving programs, assessment of environmental effects of fuel consumption, emissions reduction studies, and air traffic management applications. In this study, utilizing engine manufacturers' real data, a metaheuristic model based on genetic algorithms (GAs) and a machine learning model based on neural networks (NNs) trained with Levenberg-Marquardt (LM), delta-bar-delta (DBD), and conjugate gradient (CG) algorithms were accomplished to incorporate the effect of both flight altitude and Mach number in the estimation of thrust. For the GA model, the analysis of population size impact on the model's accuracy and effect of number of data on model coefficients were also performed. For the NN model, design of optimum topology was searched for one- and two-hidden-layer networks. Predicted thrust values presented a close agreement with real thrust data for both models, among which LM trained NNs gave the best accuracies.

Performance seeking control program overview

NASA Technical Reports Server (NTRS)

Orme, John S.

1995-01-01

The Performance Seeking Control (PSC) program evolved from a series of integrated propulsion-flight control research programs flown at NASA Dryden Flight Research Center (DFRC) on an F-15. The first of these was the Digital Electronic Engine Control (DEEC) program and provided digital engine controls suitable for integration. The DEEC and digital electronic flight control system of the NASA F-15 were ideally suited for integrated controls research. The Advanced Engine Control System (ADECS) program proved that integrated engine and aircraft control could improve overall system performance. The objective of the PSC program was to advance the technology for a fully integrated propulsion flight control system. Whereas ADECS provided single variable control for an average engine, PSC controlled multiple propulsion system variables while adapting to the measured engine performance. PSC was developed as a model-based, adaptive control algorithm and included four optimization modes: minimum fuel flow at constant thrust, minimum turbine temperature at constant thrust, maximum thrust, and minimum thrust. Subsonic and supersonic flight testing were conducted at NASA Dryden covering the four PSC optimization modes and over the full throttle range. Flight testing of the PSC algorithm, conducted in a series of five flight test phases, has been concluded at NASA Dryden covering all four of the PSC optimization modes. Over a three year period and five flight test phases 72 research flights were conducted. The primary objective of flight testing was to exercise each PSC optimization mode and quantify the resulting performance improvements.

Nuclear Cryogenic Propulsion Stage

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Borowski, S. K.; George, J. A.; Kim, T.; Emrich, W. J.; Hickman, R. R.; Broadway, J. W.; Gerrish, H. P.; Adams, R. B.

2012-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced NEP.

Wave Fluid Film Bearing Tests for an Aviation Gearbox

NASA Technical Reports Server (NTRS)

Dimofte, Florin; Proctor, Margaret P.; Fleming, David P.; Keith, Theo G., Jr.

2000-01-01

An oil-lubricated wave journal-thrust bearing assembly was successfully tested at conditions found in general aviation engine gearboxes. The bearing performed well at both steady state conditions and in start-stop tests. It ran stably under all loading conditions, including zero load, at all speeds up to 16 000 rpm. The bearing carried 25 percent more load than required for the gearbox application, supporting 8900 N (94 bars average pressure), and showed very good thermal stability. 450 start-stop cycles were also performed, including 350 cycles without oil supply during starting and stopping. Test results and numerical predictions were in good agreement.

Monopropellant hydrazine resistojet: Flight application design

NASA Technical Reports Server (NTRS)

Kurch, C. K.

1973-01-01

The design, development, and testing of an engineering model nominal 20-millipound thrust monopropellant hydrazine resistojet program is directed toward the advanced development of an electrothermal hydrazine thruster (EHT). The EHT decomposes hydrazine thermally and expands the decomposition products through a nozzle to provide the impulse necessary to fulfill spacecraft propulsive requirements. The thruster is capable of operation at pulse widths from 0.050 second to steady state and delivers specific impulse values up to about 230 seconds depending on the duty cycle. The program is comprised of six tasks including analyses, the generation of specifications and other documentation, design, fabrication and test, data correlation, and recommendations for the design of flight units.

Affordable Development of a Nuclear Cryogenic Propulsion Stage

NASA Technical Reports Server (NTRS)

Houts, M. G.; Borowski, S. K.; George, J. A.; Kim, T.; Emrich, W. J.; Hickman, R. R.; Broadway, J. W.; Gerrish, H. P.; Adams, R. B.

2012-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. The foundation provided by development and utilization of a NCPS could enable development of extremely high performance systems. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced Nuclear Electric Propulsion (NEP).

Fast Track Lunar NTR Systems Assessment for NASA's First Lunar Outpost and Its Evolvability to Mars

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Alexander, Stephen W.

1995-01-01

Integrated systems and missions studies are presented for an evolutionary lunar-to-Mars space transportation system (STS) based on nuclear thermal rocket (NTR) technology. A 'standardized' set of engine and stage components are identified and used in a 'building block' fashion to configure a variety of piloted and cargo, lunar and Mars vehicles. The reference NTR characteristics include a thrust of 50 thousand pounds force (klbf), specific impulse (I(sub sp)) of 900 seconds, and an engine thrust-to-weight ratio of 4. 3. For the National Aeronautics and Space Administrations (NASA) First Lunar Outpost (FLO) mission, and expendable NTR stage powered by two such engines can deliver approximately 96 metric tonnes (t) to trans-lunar injection (TLI) conditions for an initial mass in low Earth orbit (IMLEO) of approximately 198 t compared to 250 t for a cryogenic chemical system. The stage liquid hydrogen (LH2) tank has a diameter, length, and capacity of 10 m, 14.5 m and 66 t, respectively. By extending the stage length and LH2 capacity to approximately 20 m and 96 t, a single launch Mars cargo vehicle could deliver to an elliptical Mars parking orbit a 63 t Mars excursion vehicle (MEV) with a 45 t surface payload. Three 50 klbf engines and the two standardized LH2 tanks developed for the lunar and Mars cargo vehicles are used to configure the vehicles supporting piloted Mars missions as early as 2010. The 'modular' NTR vehicle approach forms the basis for an efficient STS able to handle the needs of a wide spectrum of lunar and Mars missions.

Research Technology

NASA Image and Video Library

2002-03-11

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Thermal Modeling for Pulsed Inductive FRC Plasmoid Thrusters

NASA Astrophysics Data System (ADS)

Pfaff, Michael

Due to the rising importance of space based infrastructure, long-range robotic space missions, and the need for active attitude control for spacecraft, research into Electric Propulsion is becoming increasingly important. Electric Propulsion (EP) systems utilize electric power to accelerate ions in order to produce thrust. Unlike traditional chemical propulsion, this means that thrust levels are relatively low. The trade-off is that EP thrusters have very high specific impulses (Isp), and can therefore make do with far less onboard propellant than cold gas, monopropellant, or bipropellant engines. As a consequence of the high power levels used to accelerate the ionized propellant, there is a mass and cost penalty in terms of solar panels and a power processing unit. Due to the large power consumption (and waste heat) from electric propulsion thrusters, accurate measurements and predictions of thermal losses are needed. Excessive heating in sensitive locations within a thruster may lead to premature failure of vital components. Between the fixed cost required to purchase these components, as well as the man-hours needed to assemble (or replace) them, attempting to build a high-power thruster without reliable thermal modeling can be expensive. This paper will explain the usage of FEM modeling and experimental tests in characterizing the ElectroMagnetic Plasmoid Thruster (EMPT) and the Electrodeless Lorentz Force (ELF) thruster at the MSNW LLC facility in Redmond, Washington. The EMPT thruster model is validated using an experimental setup, and steady state temperatures are predicted for vacuum conditions. Preliminary analysis of the ELF thruster indicates possible material failure in absence of an active cooling system for driving electronics and for certain power levels.

Feasibility of rotating fluidized bed reactor for rocket propulsion

NASA Technical Reports Server (NTRS)

Ludewig, H.; Manning, A. J.; Raseman, C. J.

1974-01-01

The rotating fluidized bed reactor concept is outlined, and its application to rocket propulsion is discussed. Experimental results obtained indicate that minimum fluidization correlations commonly in use for 1-g beds can also be applied to multiple-g beds. It was found that for a low thrust system (20,000 lbf) the fuel particle size and/or particle stress play a limiting role on performance. The superiority of U-233 as a fuel for this type of rocket engine is clearly demonstrated in the analysis. The maximum thrust/weight ratio for a 90,000N thrust engine was found to be approximately 65N/kg.

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.

Electron microscopic study of soot particulate matter emissions from aircraft turbine engines.

PubMed

Liati, Anthi; Brem, Benjamin T; Durdina, Lukas; Vögtli, Melanie; Dasilva, Yadira Arroyo Rojas; Eggenschwiler, Panayotis Dimopoulos; Wang, Jing

2014-09-16

The microscopic characteristics of soot particulate matter (PM) in gas turbine exhaust are critical for an accurate assessment of the potential impacts of the aviation industry on the environment and human health. The morphology and internal structure of soot particles emitted from a CFM 56-7B26/3 turbofan engine were analyzed in an electron microscopic study, down to the nanoscale, for ∼ 100%, ∼ 65%, and ∼ 7% static engine thrust as a proxy for takeoff, cruising, and taxiing, respectively. Sampling was performed directly on transmission electron microscopy (TEM) grids with a state-of-the-art sampling system designed for nonvolatile particulate matter. The electron microscopy results reveal that ∼ 100% thrust produces the highest amount of soot, the highest soot particle volume, and the largest and most crystalline primary soot particles with the lowest oxidative reactivity. The opposite is the case for soot produced during taxiing, where primary soot particles are smallest and most reactive and the soot amount and volume are lowest. The microscopic characteristics of cruising condition soot resemble the ones of the ∼ 100% thrust conditions, but they are more moderate. Real time online measurements of number and mass concentration show also a clear correlation with engine thrust level, comparable with the TEM study. The results of the present work, in particular the small size of primary soot particles present in the exhaust (modes of 24, 20, and 13 nm in diameter for ∼ 100%, ∼ 65% and ∼ 7% engine thrust, respectively) could be a concern for human health and the environment and merit further study. This work further emphasizes the significance of the detailed morphological characteristics of soot for assessing environmental impacts.

Modular thrust subsystem approaches to solar electric propulsion module design

NASA Technical Reports Server (NTRS)

Cake, J. E.; Sharp, G. R.; Oglebay, J. C.; Shaker, F. J.; Zavesky, R. J.

1976-01-01

Three approaches are presented for packaging the elements of a 30 cm ion thruster subsystem into a modular thrust subsystem. The individual modules, when integrated into a conceptual solar electric propulsion module are applicable to a multimission set of interplanetary flights with the space shuttle interim upper stage as the launch vehicle. The emphasis is on the structural and thermal integration of the components into the modular thrust subsystems. Thermal control for the power processing units is either by direct radiation through louvers in combination with heat pipes or an all heat pipe system. The propellant storage and feed system and thruster gimbal system concepts are presented. The three approaches are compared on the basis of mass, cost, testing, interfaces, simplicity, reliability, and maintainability.

Modular thrust subsystem approaches to solar electric propulsion module design

NASA Technical Reports Server (NTRS)

Cake, J. E.; Sharp, G. R.; Oglebay, J. C.; Shaker, F. J.; Zevesky, R. J.

1976-01-01

Three approaches are presented for packaging the elements of a 30 cm ion thrustor subsystem into a modular thrust subsystem. The individual modules, when integrated into a conceptual solar electric propulsion module are applicable to a multimission set of interplanetary flights with the Space Shuttle/Interim Upper Stage as the launch vehicle. The emphasis is on the structural and thermal integration of the components into the modular thrust subsystems. Thermal control for the power processing units is either by direct radiation through louvers in combination with heat pipes of an all heat pipe system. The propellant storage and feed system and thrustor gimbal system concepts are presented. The three approaches are compared on the basis of mass, cost, testing, interfaces, simplicity, reliability, and maintainability.

Characterization of Gas-Phase Organics Using Proton Transfer Reaction Time-of-Flight Mass Spectrometry: Aircraft Turbine Engines.

PubMed

Kilic, Dogushan; Brem, Benjamin T; Klein, Felix; El-Haddad, Imad; Durdina, Lukas; Rindlisbacher, Theo; Setyan, Ari; Huang, Rujin; Wang, Jing; Slowik, Jay G; Baltensperger, Urs; Prevot, Andre S H

2017-04-04

Nonmethane organic gas emissions (NMOGs) from in-service aircraft turbine engines were investigated using a proton transfer reaction time-of-flight mass spectrometer (PTR-ToF-MS) at an engine test facility at Zurich Airport, Switzerland. Experiments consisted of 60 exhaust samples for seven engine types (used in commercial aviation) from two manufacturers at thrust levels ranging from idle to takeoff. Emission indices (EIs) for more than 200 NMOGs were quantified, and the functional group fractions (including acids, carbonyls, aromatics, and aliphatics) were calculated to characterize the exhaust chemical composition at different engine operation modes. Total NMOG emissions were highest at idling with an average EI of 7.8 g/kg fuel and were a factor of ∼40 lower at takeoff thrust. The relative contribution of pure hydrocarbons (particularly aromatics and aliphatics) of the engine exhaust decreased with increasing thrust while the fraction of oxidized compounds, for example, acids and carbonyls increased. Exhaust chemical composition at idle was also affected by engine technology. Older engines emitted a higher fraction of nonoxidized NMOGs compared to newer ones. Idling conditions dominated ground level organic gas emissions. Based on the EI determined here, we estimate that reducing idle emissions could substantially improve air quality near airports.

Real-time in-flight engine performance and health monitoring techniques for flight research application

NASA Technical Reports Server (NTRS)

Ray, Ronald J.; Hicks, John W.; Wichman, Keith D.

1991-01-01

Procedures for real time evaluation of the inflight health and performance of gas turbine engines and related systems were developed to enhance flight test safety and productivity. These techniques include the monitoring of the engine, the engine control system, thrust vectoring control system health, and the detection of engine stalls. Real time performance techniques were developed for the determination and display of inflight thrust and for aeroperformance drag polars. These new methods were successfully shown on various research aircraft at NASA-Dryden. The capability of NASA's Western Aeronautical Test Range and the advanced data acquisition systems were key factors for implementation and real time display of these methods.

Internal performance predictions for Langley scramjet engine module

NASA Technical Reports Server (NTRS)

Pinckney, S. Z.

1978-01-01

A one dimensional theoretical method for the prediction of the internal performance of a scramjet engine is presented. The effects of changes in vehicle forebody flow parameters and characteristics on predicted thrust for the scramjet engine were evaluated using this method, and results are presented. A theoretical evaluation of the effects of changes in the scramjet engine's internal parameters is also presented. Theoretical internal performance predictions, in terms thrust coefficient and specific impulse, are provided for the scramjet engine for free stream Mach numbers of 5, 6, and 7 free stream dynamic pressure of 23,940 N/sq m forebody surface angles of 4.6 deg to 14.6 deg, and fuel equivalence ratio of 1.0.

Status of the NEXT Ion Engine Wear Test

NASA Technical Reports Server (NTRS)

Soulas, George C.; Domonkos, Matthew T.; Kamhawi, Hani; Patterson, Michael J.; Gardner, Michael M.

2003-01-01

The status of the NEXT 2000 hour wear test is presented. This test is being conducted with a 40 cm engineering model ion engine, designated EM1, at a beam current higher than listed on the NEXT throttle table. Pretest performance assessments demonstrated that EM1 satisfies all thruster performance requirements. As of 7/3/03, the ion engine has accumulated 406 hours of operation at a thruster input power of 6.9 kW. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with no indications of performance degradation. Images of the downstream discharge cathode, neutralizer, and accelerator aperture surfaces have exhibited no significant erosion to date.

Control technology for future aircraft propulsion systems

NASA Technical Reports Server (NTRS)

Zeller, J. R.; Szuch, J. R.; Merrill, W. C.; Lehtinen, B.; Soeder, J. F.

1984-01-01

The need for a more sophisticated engine control system is discussed. The improvements in better thrust-to-weight ratios demand the manipulation of more control inputs. New technological solutions to the engine control problem are practiced. The digital electronic engine control (DEEC) system is a step in the evolution to digital electronic engine control. Technology issues are addressed to ensure a growth in confidence in sophisticated electronic controls for aircraft turbine engines. The need of a control system architecture which permits propulsion controls to be functionally integrated with other aircraft systems is established. Areas of technology studied include: (1) control design methodology; (2) improved modeling and simulation methods; and (3) implementation technologies. Objectives, results and future thrusts are summarized.

Investigation of Water-spray Cooling of Turbine Blades in a Turbojet Engine

NASA Technical Reports Server (NTRS)

Freche, John C; Stelpflug, William J

1953-01-01

An analytical and experimental investigation was made with a J33-A-9 engine to determine the effectiveness of spray cooling as a means of increasing thrust by permitting engine operation at inlet-gas temperatures and speeds above rated. With the assumption of adequate spray cooling at a coolant-to-gas flow ratio of 3 percent, calculations for the sea-level static condition indicated a thrust may be achieved by engine operation at an inlet-gas temperature of 2000 degrees F and an overspeed of 10 percent. Of the water-injection configurations investigated experimentally, those located in the inner ring of the stator diaphragm provided the best cooling at rated engine speed.

A New Reliability Analysis Model of the Chegongzhuang Heat-Supplying Tunnel Structure Considering the Coupling of Pipeline Thrust and Thermal Effect

PubMed Central

Zhang, Jiawen; He, Shaohui; Wang, Dahai; Liu, Yangpeng; Yao, Wenbo; Liu, Xiabing

2018-01-01

Based on the operating Chegongzhuang heat-supplying tunnel in Beijing, the reliability of its lining structure under the action of large thrust and thermal effect is studied. According to the characteristics of a heat-supplying tunnel service, a three-dimensional numerical analysis model was established based on the mechanical tests on the in-situ specimens. The stress and strain of the tunnel structure were obtained before and after the operation. Compared with the field monitoring data, the rationality of the model was verified. After extracting the internal force of the lining structure, the improved method of subset simulation was proposed as the performance function to calculate the reliability of the main control section of the tunnel. In contrast to the traditional calculation method, the analytic relationship between the sample numbers in the subset simulation method and Monte Carlo method was given. The results indicate that the lining structure is greatly influenced by coupling in the range of six meters from the fixed brackets, especially the tunnel floor. The improved subset simulation method can greatly save computation time and improve computational efficiency under the premise of ensuring the accuracy of calculation. It is suitable for the reliability calculation of tunnel engineering, because “the lower the probability, the more efficient the calculation.” PMID:29401691

Space Launch Initiative (SLI) Engine Test

NASA Technical Reports Server (NTRS)

2002-01-01

NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama, has begun a series of engine tests on the Reaction Control Engine developed by TRW Space and Electronics for NASA's Space Launch Initiative (SLI). SLI is a technology development effort aimed at improving the safety, reliability, and cost effectiveness of space travel for reusable launch vehicles. The engine in this photo, the first engine tested at MSFC that includes SLI technology, was tested for two seconds at a chamber pressure of 185 pounds per square inch absolute (psia). Propellants used were liquid oxygen as an oxidizer and liquid hydrogen as fuel. Designed to maneuver vehicles in orbit, the engine is used as an auxiliary propulsion system for docking, reentry, fine-pointing, and orbit transfer while the vehicle is in orbit. The Reaction Control Engine has two unique features. It uses nontoxic chemicals as propellants, which creates a safer environment with less maintenance and quicker turnaround time between missions, and it operates in dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The force of low level thrust allows the vehicle to fine-point maneuver and dock, while the force of the high level thrust is used for reentry, orbital transfer, and course positioning.

Test and evaluation of the HIDEC engine uptrim algorithm. [Highly Integrated Digital Electronic Control for aircraft

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemente into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

Reverse Flow Engine Core Having a Ducted Fan with Integrated Secondary Flow Blades

NASA Technical Reports Server (NTRS)

Kisska, Michael K. (Inventor); Princen, Norman H. (Inventor); Kuehn, Mark S. (Inventor); Cosentino, Gary B. (Inventor)

2014-01-01

Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Technical Excellence: A Requirement for Good Engineering

NASA Technical Reports Server (NTRS)

Gill, Paul S.; Vaughan, William W.

2008-01-01

Technical excellence is a requirement for good engineering. Technical excellence has many different ways of expressing itself within engineering. NASA has initiatives that address the enhancement of the Agency's technical excellence and thrust to maintain the associated high level of performance by the Agency on current programs/projects and as it moves into the Constellation Program and the return to the Moon with plans to visit Mars. This paper addresses some of the key initiatives associated with NASA's technical excellence thrust. Examples are provided to illustrate some results being achieved and plans to enhance these initiatives.

Orbit transfer vehicle engine study, phase A, extension 1: Volume 2: Study results

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Because of the advantage of the Advanced Expander Cycle Engine brought out in initial studies, further design optimization and comparative analyses were undertaken. The major results and conclusion derived are summarized. The primary areas covered are (1) thrust chamber geometry optimization, (2) expander cycle optimization, (3) alternate low thrust capability, (4) safety and reliability, (5) development risk comparison, and (6) cost comparisons. All of the results obtained were used to baseline the initial design concept for the OTV Advanced Expander Cycle Engine Point Design Study.

Engine sizing and integration requirements for hypersonic airbreathing missile applications

NASA Astrophysics Data System (ADS)

Waltrup, P. J.; Billig, F. S.; Stockbridge, R. D.

1982-03-01

A procedure that provides a rational means for selecting an inlet/combustor configuration for a hypersonic airbreathing missile is presented. The particular problem that is addressed is the design of the sustained engine of a two stage missile that is constrained to be launched from a stowage volume that is either square or circular in cross section. The sustainer engine accelerates from a low altitude separation at Mach 4 and climbs to high altitude for cruise at Mach 8. The results show that a missile with an axisymmetric nose inlet provides a somewhat higher thrust capability and slightly better fuel efficiency than a chin type inlet. Aft entry inlets are shown to have a substantially lower thrust potential and lower engine efficiency. A criterion for determining the maximum contraction ratio of a fixed geometry inlet is established and applied to the exemplary missile designs. Combustor area ratio is examined and found to have a relatively small effect on engine performance for area ratios equal to or larger than that required to obtain maximum thrust at the take-over Mach number.

Oxygen-hydrogen thrusters for Space Station auxiliary propulsion systems

NASA Technical Reports Server (NTRS)

Berkman, D. K.

1984-01-01

The feasibility and technology requirements of a low-thrust, high-performance, long-life, gaseous oxygen (GO2)/gaseous hydrogen (GH2) thruster were examined. Candidate engine concepts for auxiliary propulsion systems for space station applications were identified. The low-thrust engine (5 to 100 lb sub f) requires significant departure from current applications of oxygen/hydrogen propulsion technology. Selection of the thrust chamber material and cooling method needed or long life poses a major challenge. The use of a chamber material requiring a minimum amount of cooling or the incorporation of regenerative cooling were the only choices available with the potential of achieving very high performance. The design selection for the injector/igniter, the design and fabrication of a regeneratively cooled copper chamber, and the design of a high-temperature rhenium chamber were documented and the performance and heat transfer results obtained from the test program conducted at JPL using the above engine components presented. Approximately 115 engine firings were conducted in the JPL vacuum test facility, using 100:1 expansion ratio nozzles. Engine mixture ratio and fuel-film cooling percentages were parametrically investigated for each test configuration.

Thrust control system design of ducted rockets

NASA Astrophysics Data System (ADS)

Chang, Juntao; Li, Bin; Bao, Wen; Niu, Wenyu; Yu, Daren

2011-07-01

The investigation of the thrust control system is aroused by the need for propulsion system of ducted rockets. Firstly the dynamic mathematical models of gas flow regulating system, pneumatic servo system and ducted rocket engine were established and analyzed. Then, to conquer the discussed problems of thrust control, the idea of information fusion was proposed to construct a new feedback variable. With this fused feedback variable, the thrust control system was designed. According to the simulation results, the introduction of the new fused feedback variable is valid in eliminating the contradiction between rapid response and stability for the thrust control system of ducted rockets.

14 CFR Appendix B to Part 121 - Airplane Flight Recorder Specification

Code of Federal Regulations, 2010 CFR

2010-01-01

... Transmitter Keying On-Off (Discrete) ±2° ±2% Thrust/Power on Each Engine Full Range Forward ±2° 1 (per engine) 0.2% 2 Trailing Edge Flap or Cockpit Control Selection Full Range or Each Discrete Position ±3° or... Discrete Position ±3° or as Pilot's Indicator 0.5 0.5% 2 Thrust Reverser Position Stowed, In Transit, and...

Orbital transfer vehicle engine technology: Baffled injector design, fabrication, and verification

NASA Technical Reports Server (NTRS)

Schneider, J. A.

1991-01-01

New technologies for space-based, reusable, throttleable, cryogenic orbit transfer propulsion are being evaluated. Supporting tasks for the design of a dual expander cycle engine thrust chamber design are documented. The purpose of the studies was to research the materials used in the thrust chamber design, the supporting fabrication methods necessary to complete the design, and the modification of the injector element for optimum injector/chamber compatibility.

Twin jet shielding. [for aircraft noise reduction

NASA Technical Reports Server (NTRS)

Parthasarathy, S. P.; Cuffel, R. F.; Massier, P. F.

1979-01-01

For an over-the-wing/under-the-wing engine configuration on an airplane, the noise produced by the upper jet flow is partially reflected by the lower jet. An analysis has been performed which can be used to predict the distribution of perceived noise levels along the ground plane at take-off for an airplane which is designed to take advantage of the over/under shielding concept. Typical contours of PNL, the shielding benefit in the shadow zone, and the EPNL values at 3.5 nautical miles from brake release as well as EPNL values at sideline at 0.35 nautical miles have been calculated. This has been done for a range of flow parameters characteristic of engines producing inverted velocity profile jets suitable for use in a supersonic cruise vehicle. Reductions up to 6.0 EPNdB in community noise levels can be realized when the over engines are operated at higher thrust and the lower engines simultaneously operated with reduced thrust keeping the total thrust constant.

Tests and analysis of a vented D thrust deflecting nozzle on a turbofan engine. [conducted at the outdoor aerodynamic research facility of the Ames Research Center

NASA Technical Reports Server (NTRS)

Roseberg, E. W.

1982-01-01

The objectives were to: obtain nozzle performance characteristics in and out of ground effects; demonstrate the compatibility of the nozzle with a turbofan engine; obtain pressure and temperature distributions on the surface of the D vented nozzle; and establish a correlation of the nozzle performance between small scale and large scale models. The test nozzle was a boilerplate model of the MCAIR D vented nozzle configured for operation with a General Electric YTF-34-F5 turbofan engine. The nozzle was configured to provide: a thrust vectoring range of 0 to 115 deg; a yaw vectoring range of 0 to 10 deg; variable nozzle area control; and variable spacing between the core exit and nozzle entrance station. Compatibility between the YTF-34-T5 turbofan engine and the D vented nozzle was demonstrated. Velocity coefficients of 0.96 and greater were obtained for 90 deg of thrust vectoring. The nozzle walls remained cool during all test conditions.

Numerical Simulations of Single Flow Element in a Nuclear Thermal Thrust Chamber

NASA Technical Reports Server (NTRS)

Cheng, Gary; Ito, Yasushi; Ross, Doug; Chen, Yen-Sen; Wang, Ten-See

2007-01-01

The objective of this effort is to develop an efficient and accurate computational methodology to predict both detailed and global thermo-fluid environments of a single now element in a hypothetical solid-core nuclear thermal thrust chamber assembly, Several numerical and multi-physics thermo-fluid models, such as chemical reactions, turbulence, conjugate heat transfer, porosity, and power generation, were incorporated into an unstructured-grid, pressure-based computational fluid dynamics solver. The numerical simulations of a single now element provide a detailed thermo-fluid environment for thermal stress estimation and insight for possible occurrence of mid-section corrosion. In addition, detailed conjugate heat transfer simulations were employed to develop the porosity models for efficient pressure drop and thermal load calculations.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

Workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, hoist the thrust structure assembly for the Saturn IB S-IB (first) stage. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Ignition Characterization Test Results for the LO2/Ethanol Propellant Combination

NASA Technical Reports Server (NTRS)

Popp, Christopher G.; Robinson, Phillip J.; Veith, Eric M.

2006-01-01

A series of contracts were issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) under the auspices of the Exploration Systems Mission Directorate to develop and expand the maturity of candidate technologies considered to be important for future space exploration. One such technology was to determine the viability of incorporating non-toxic propellants for Reaction Control Subsystems (RCS). Contract NAS8-01109 was issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilized liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporated a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. The preferred RCS approach for the dual thrust RCE was to utilize pressure-fed liquid oxygen (LOX) and ethanol propellants; however, previous dual thrust feasibility testing incorporated GOX/Ethanol igniters as opposed to LOX/Ethanol igniters in the design. GOX/Ethanol was easier to ignite, but this combination had system design implications of providing GOX for the igniters. A LOX/Ethanol igniter was desired; however, extensive LOX/Ethanol ignition data over the anticipated operating range for the dual thrust RCE did not exist. Therefore, Aerojet designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk mitigation effort for the dual thrust RCE design. LOX, encompassing potential two-phase flow conditions anticipated being present in real mission applications. A workhorse igniter was designed to accommodate the hll LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the latent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to subcooled LOX. permitted oxygen state points to be determined in the igniter oxidizer manifold, and gas-side igniter chamber thermocouples provided chamber thermal profile characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter (P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, m. Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 210 R LOX to 486 R GOX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5. Pulsing is required in typical RCS engines; therefore, the workhorse igniter was pulse tested to verify the ability to provide the required ignition for a pulsing RCE. The minimum electrical pulse width (EPW) of the dual thrust RCE was 0.080 seconds.

An inventory of aeronautical ground research facilities. Volume 2: Air breathing engine test facilities

NASA Technical Reports Server (NTRS)

Pirrello, C. J.; Hardin, R. D.; Heckart, M. V.; Brown, K. R.

1971-01-01

The inventory covers free jet and direct connect altitude cells, sea level static thrust stands, sea level test cells with ram air, and propulsion wind tunnels. Free jet altitude cells and propulsion wind tunnels are used for evaluation of complete inlet-engine-exhaust nozzle propulsion systems under simulated flight conditions. These facilities are similar in principal of operation and differ primarily in test section concept. The propulsion wind tunnel provides a closed test section and restrains the flow around the test specimen while the free jet is allowed to expand freely. A chamber of large diameter about the free jet is provided in which desired operating pressure levels may be maintained. Sea level test cells with ram air provide controlled, conditioned air directly to the engine face for performance evaluation at low altitude flight conditions. Direct connect altitude cells provide a means of performance evaluation at simulated conditions of Mach number and altitude with air supplied to the flight altitude conditions. Sea level static thrust stands simply provide an instrumented engine mounting for measuring thrust at zero airspeed. While all of these facilities are used for integrated engine testing, a few provide engine component test capability.

Design, analysis, and control of a large transport aircraft utilizing selective engine thrust as a backup system for the primary flight control. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Gerren, Donna S.

1995-01-01

A study has been conducted to determine the capability to control a very large transport airplane with engine thrust. This study consisted of the design of an 800-passenger airplane with a range of 5000 nautical miles design and evaluation of a flight control system, and design and piloted simulation evaluation of a thrust-only backup flight control system. Location of the four wing-mounted engines was varied to optimize the propulsive control capability, and the time constant of the engine response was studied. The goal was to provide level 1 flying qualities. The engine location and engine time constant did not have a large effect on the control capability. The airplane design did meet level 1 flying qualities based on frequencies, damping ratios, and time constants in the longitudinal and lateral-directional modes. Project pilots consistently rated the flying qualities as either level 1 or level 2 based on Cooper-Harper ratings. However, because of the limited control forces and moments, the airplane design fell short of meeting the time required to achieve a 30 deg bank and the time required to respond a control input.

Assessment of community noise for a medium-range airplane with open-rotor engines

NASA Astrophysics Data System (ADS)

Kopiev, V. F.; Shur, M. L.; Travin, A. K.; Belyaev, I. V.; Zamtfort, B. S.; Medvedev, Yu. V.

2017-11-01

Community noise of a hypothetical medium-range airplane equipped with open-rotor engines is assessed by numerical modeling of the aeroacoustic characteristics of an isolated open rotor with the simplest blade geometry. Various open-rotor configurations are considered at constant thrust, and the lowest-noise configuration is selected. A two-engine medium-range airplane at known thrust of bypass turbofan engines at different segments of the takeoff-landing trajectory is considered, after the replacement of those engines by the open-rotor engines. It is established that a medium-range airplane with two open-rotor engines meets the requirements of Chapter 4 of the ICAO standard with a significant margin. It is shown that airframe noise makes a significant contribution to the total noise of an airplane with open-rotor engines at landing.

Experimental study of the thermal-acoustic efficiency in a long turbulent diffusion-flame burner

NASA Technical Reports Server (NTRS)

Mahan, J. R.

1983-01-01

A two-year study of noise production in a long tubular burner is described. The research was motivated by an interest in understanding and eventually reducing core noise in gas turbine engines. The general approach is to employ an acoustic source/propagation model to interpret the sound pressure spectrum in the acoustic far field of the burner in terms of the source spectrum that must have produced it. In the model the sources are assumed to be due uniquely to the unsteady component of combustion heat release; thus only direct combustion-noise is considered. The source spectrum is then the variation with frequency of the thermal-acoustic efficiency, defined as the fraction of combustion heat release which is converted into acoustic energy at a given frequency. The thrust of the research was to study the variation of the source spectrum with the design and operating parameters of the burner.