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At least 883 records · Page 49
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Turbomachine Sealing and Secondary Flows - Part 3: Review of Power-Stream Support, Unsteady Flow Systems, Seal and Disk Cavity Flows, Engine Externals, and Life and Reliability Issues - Part 3

The issues and components supporting the engine power stream are reviewed. It is essential that companies pay close attention to engine sealing issues, particularly on the high-pressure spool or high-pressure pumps. Small changes in these systems are reflected throughout the entire engine. Although cavity, platform, and tip sealing are complex and have a significant effect on component and engine performance, computational tools (e.g., NASA-developed INDSEAL, SCISEAL, and ADPAC) are available to help guide the designer and the experimenter. Gas turbine engine and rocket engine externals must all function efficiently with a high degree of reliability in order for the engine to run but often receive little attention until they malfunction. Within the open literature statistically significant data for critical engine components are virtually nonexistent; the classic approach is deterministic. Studies show that variations with loading can have a significant effect on component performance and life. Without validation data they are just studies. These variations and deficits in statistical databases require immediate attention.

Hendricks, R. C.↗
NASA NTRS1948

Altitude-Wind-Tunnel Investigation of Westinghouse 19B-2 19B-8, and 19XB-1 Jet-Propulsion Engines: Operational Characteristics - Part 1

An investigation was conducted in the NACA Cleveland altitude wind tunnel to determine the operational characteristics of the Westinghouse 19B-2, 19B-8, and 19XB-l jet-propulsion engines. The 19B engine is one af the earliest experimental Westinghouse axial flow engines. The 19XB-1 engine is an experimental prototype of the Westinghouse 15 series, having a rated thrust of 1400 pounds. Improvements in performance and operational characteristics have resulted in the 19XB-2B engine with a rated thrust of 1600 pounds. The operational characteristics were determined over a range of simulated altitudes from 5000 to 30,000 feet for the 19B engines and from 5000 to 35000 feet for the 19XB-l engine at airspeed from 20 to 380 miles per hour. The affects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, starting, acceleration, and functioning of the fuel-control system are discussed. Damage to the engines that occurred during the investigation is also briefly discussed. The changes made in the combustion-chamber configuration to improve the operating we are described.

Fleming, William A.↗
NASA NTRS2005

Intelligent Life-Extending Controls for Aircraft Engines

Aircraft engine controllers are designed and operated to provide desired performance and stability margins. The purpose of life-extending-control (LEC) is to study the relationship between control action and engine component life usage, and to design an intelligent control algorithm to provide proper trade-offs between performance and engine life usage. The benefit of this approach is that it is expected to maintain safety while minimizing the overall operating costs. With the advances of computer technology, engine operation models, and damage physics, it is necessary to reevaluate the control strategy fro overall operating cost consideration. This paper uses the thermo-mechanical fatigue (TMF) of a critical component to demonstrate how an intelligent engine control algorithm can drastically reduce the engine life usage with minimum sacrifice in performance. A Monte Carlo simulation is also performed to evaluate the likely engine damage accumulation under various operating conditions. The simulation results show that an optimized acceleration schedule can provide a significant life saving in selected engine components.

Guo, Ten-Huei↗
NASA NTRS1997

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

Source record↗
NASA NTRS2004

Probabilistic Study Conducted on Sensor-Based Engine Life Calculation

Turbine engine life management is a very complicated process to ensure the safe operation of an engine subjected to complex usage. The challenge of life management is to find a reasonable compromise between the safe operation and the maximum usage of critical parts to reduce maintenance costs. The commonly used "cycle count" approach does not take the engine operation conditions into account, and it oversimplifies the calculation of the life usage. Because of the shortcomings, many engine components are regularly pulled for maintenance before their usable life is over. And, if an engine has been running regularly under more severe conditions, components might not be taken out of service before they exceed their designed risk of failure. The NASA Glenn Research Center and its industrial and academic partners have been using measurable parameters to improve engine life estimation. This study was based on the Monte Carlo simulation of 5000 typical flights under various operating conditions. First a closed-loop engine model was developed to simulate the engine operation across the mission profile and a thermomechanical fatigue (TMF) damage model was used to calculate the actual damage during takeoff, where the maximum TMF accumulates. Next, a Weibull distribution was used to estimate the implied probability of failure for a given accumulated cycle count. Monte Carlo simulations were then employed to find the profiles of the TMF damage under different operating assumptions including parameter uncertainties. Finally, probabilities of failure for different operating conditions were analyzed to demonstrate the importance of a sensor-based damage calculation in order to better manage the risk of failure and on-wing life.

Guo, Ten-Huei↗
NASA NTRS2001

Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high-cycle thermal fatigue behavior has been investigated on a flat Haynes 188 alloy specimen, under the test condition of 30-Hz cycle frequency (33-msec pulse period and 10-msec pulse width including a 0.2-msec pulse spike; ref. 4). Temperature distributions were calculated with one-dimensional finite difference models. The calculations show that that the 0.2-msec pulse spike can cause an additional 40 C temperature fluctuation with an interaction depth of 0.08 mm near the specimen surface region. This temperature swing will be superimposed onto the temperature swing of 80 C that is induced by the 10-msec laser pulse near the 0.53-mm-deep surface interaction region.

Zhu, Dong-Ming↗
NASA NTRS2006

Structural Requirements for the Space Propulsion Engine Systems

In January 2004, the National Aeronautics and Space Administration (NASA) was given a vision for Space Exploration by President Bush, setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. As NASA gets ready to meet the vision set by President Bush, failures are not an option. Reliability of the propulsion engine systems will play an important role in establishing an overall safe and reliable operation of these new space systems. A new standard, NASA-STD-5012, Strength and Life Assessment for Space Propulsion System Engines, has been developed to provide structural requirements for assessment of the propulsion systems engine. This standard is a complement to the current NASA-wide standard NASA-STD-5001, Structural Design and Test Factors of Safety for Spaceflight Hardware, which excluded the requirement for the engine systems (rotatory structures) along with pressure vessels. As developed, this document builds on the heritage of the multiple industrial standards related to strength and life assessment of the structures. For assuring a safe and reliable operation of a product and/or mission, establishing a set of structural assessment requirements is a key ingredient. Hence, a concentrated effort was made to improve the requirements where there are known lessons learned during the design, test, and operation phases of the Space Shuttle Main Engine (SSME) and other engine development programs. Requirements delineated in this standard are also applicable for the reusable and/or human missions. It shall be noted that "reliability of a system cannot be tested and inspected but can only be achieved if it is first designed into a system." Hence, these strength and life assessment requirements for the space propulsion system engines shall be used along with other good engineering practices, requirements, and policies.

Aggarwal, Pravin K.↗
NASA NTRS2006

Integrating International Engineering Organizations For Successful ISS Operations

The International Space Station (ISS) is a multinational orbiting space laboratory that is built in cooperation with 16 nations. The design and sustaining engineering expertise is spread worldwide. As the number of Partners with orbiting elements on the ISS grows, the challenge NASA is facing as the ISS integrator is to ensure that engineering expertise and data are accessible in a timely fashion to ensure ongoing operations and mission success. Integrating international engineering teams requires definition and agreement on common processes and responsibilities, joint training and the emergence of a unique engineering team culture. ISS engineers face daunting logistical and political challenges regarding data sharing requirements. To assure systematic information sharing and anomaly resolution of integrated anomalies, the ISS Partners are developing multi-lateral engineering interface procedures. Data sharing and individual responsibility are key aspects of this plan. This paper describes several examples of successful multilateral anomaly resolution. These successes were used to form the framework of the Partner to Partner engineering interface procedures, and this paper describes those currently documented multilateral engineering processes. Furthermore, it addresses the challenges experienced to date, and the forward work expected in establishing a successful working relationship with Partners as their hardware is launched.

Blome, Elizabeth↗
NASA NTRS2006

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

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

Snoddy, Jimmy R.↗
NASA NTRS2006

Heat Transfer and Fluid Dynamics Measurements in the Expansion Space of a Stirling Cycle Engine

The heater (or acceptor) of a Stirling engine, where most of the thermal energy is accepted into the engine by heat transfer, is the hottest part of the engine. Almost as hot is the adjacent expansion space of the engine. In the expansion space, the flow is oscillatory, impinging on a two-dimensional concavely-curved surface. Knowing the heat transfer on the inside surface of the engine head is critical to the engine design for efficiency and reliability. However, the flow in this region is not well understood and support is required to develop the CFD codes needed to design modern Stirling engines of high efficiency and power output. The present project is to experimentally investigate the flow and heat transfer in the heater head region. Flow fields and heat transfer coefficients are measured to characterize the oscillatory flow as well as to supply experimental validation for the CFD Stirling engine design codes. Presented also is a discussion of how these results might be used for heater head and acceptor region design calculations.

Jiang, Nan↗
NASA NTRS2007

Performance Optimization of Storable Bipropellant Engines to Fully Exploit Advanced Material Technologies

This paper summarizes the work performed to dale on the NASA Cycle 3A Advanced Chemical Propulsion Technology Program. The primary goals of the program are to design, fabricate, and test high performance bipropellant engines using iridium/rhenium chamber technology to obtain 335 seconds specific impulse with nitrogen tetroxide/hydrazine propellants and 330 seconds specific impulse with nitrogen tetroxide/monomethylhydrazine propellants. Aerojet has successfully completed the Base Period of this program, wherein (1) mission and system studies have been performed to verify system performance benefits and to determine engine physical and operating parameters, (2) preliminary chamber and nozzle designs have been completed and a chamber supplier has been downselected, (3) high temperature, high pressure off-nominal hot fire testing of an existing state-of-the-art high performance bipropellant engine has been completed, and (4) thermal and performance data from the engine test have been correlated with new thermal models to enable design of the new engine injector and injector/chamber interface. In the next phase of the program, Aerojet will complete design, fabrication, and test of the nitrogen tetroxide/hydrazine engine to demonstrate 335 seconds specific impulse, and also investigate improved technologies for iridium/rhenium chamber fabrication. Achievement of the NRA goals will significantly benefit NASA interplanetary missions and other government and commercial opportunities by enabling reduced launch weight and/or increased payload. At the conclusion of the program, the objective is to have an engine ready for final design and qualification for a specific science mission or commercial application. The program also constitutes a stepping stone to future, development, such as higher pressure pump-fed in-space storable engines.

Miller, Scott↗
NASA NTRS2007

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

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.

Litt, Jonathan S.↗
NASA NTRS2006

Overcoming Present-Day Powerplant Limitations Via Unconventional Engine Configurations

The Army Research Laboratory s Vehicle Technology Directorate is sponsoring the prototype development of three unconventional engine concepts - two intermittent combustion (IC) engines and one turbine engine (via SBIR (Small Business Innovative Research) contracts). The IC concepts are the Nutating Engine and the Bonner Engine, and the turbine concept is the POWER Engine. Each of the three engines offers unique and greatly improved capabilities (which cannot be achieved by present-day powerplants), while offering significant reductions in size and weight. This paper presents brief descriptions of the physical characteristics of the three engines, and discusses their performance potentials, as well as their development status.

Meitner, Peter L.↗
NASA NTRS2008

System Engineering for J-2X Development: The Simpler, the Better

The Ares I and Ares V Vehicles will utilize the J-2X rocket engine developed for NASA by the Pratt and Whitney Rocketdyne Company (PWR) as the upper stage engine (USE). The J-2X is an improved higher power version of the original J-2 engine used for Apollo. System Engineering (SE) facilitates direct and open discussions of issues and problems. This simple idea is often overlooked in large, complex engineering development programs. Definition and distribution of requirements from the engine level to the component level is controlled by Allocation Reports which breaks down numerical design objectives (weight, reliability, etc.) into quanta goals for each component area. Linked databases of design and verification requirements help eliminate redundancy and potential mistakes inherent in separated systems. Another tool, the Architecture Design Description (ADD), is used to control J-2X system architecture and effectively communicate configuration changes to those involved in the design process. But the proof of an effective process is in successful program accomplishment. SE is the methodology being used to meet the challenge of completing J-2X engine certification 2 years ahead of any engine program ever developed at PWR. This paper describes the simple, better SE tools and techniques used to achieve this success.

Kelly, William M.↗
NASA NTRS1950

Investigation of Altitude Ignition, Acceleration, and Steady-State Operation with a Single Combustor of a J47 Turbojet Engine

An investigation was conducted with a single combustor from a J47 turbojet engine using weathered aviation gasoline and several spark-plug modifications to determine altitude ignition, acceleration, and steady state operating characteristics. Satisfactory ignition was obtained with two modifications of the original opposite-polarity spark plug up to and including an altitude of 40,003 feet at conditions simulating equilibrium windmilling of the engine at a flight speed of 400 miles per hour. At a simulated altitude of 30,000 feet, satisfactory ignition was obtained over a range of simulated engine speeds. No significant effect of fuel temperature on ignition limits was observed over a range of fuel temperatures from 80 deg to -52 deg F. At an altitude of 30,000 feet, the excess temperature rise available for acceleration at low engine speeds was limited by the ability of the combustor to produce temperature rise, whereas at high engine speeds the maximum allowable turbine-inlet temperature became the restricting factor. Altitude operational limits increased from about 51,500 feet at 55 percent of rated engine speed to about 64,500 feet at 85 percent of rated speed. Combustion efficiencies varied from 59.0 to 92.6 percent over the range investigated and decreased with a decrease in engine speed and with an increase in altitude; higher efficiencies would have been obtained if lower altitudes had been investigated. Comparisons were made of the combustion efficiencies of weathered aviation gasoline and MIL-F-5616 fuel at altitudes of 30,000 and 40,000 feet. Combustion efficiencies obtained with MIL-F-5616 fuel were 8 percent higher at rated engine speed and 14 percent lower at 55 percent of rated speed than those obtained with weathered aviation gasoline.

Cook, William P.↗
NASA NTRS1954

Preliminary Altitude Operational Characteristics of a J57-P1 Turbojet Engine

The operational characteristics of a J57-P1 turbojet engine have been investigated at altitudes between 15,000 and 66,000 feet in the Lewis altitude wind tunnel. Included in this study is a discussion of fuel nozzle coking, the altitude operating limits with and without the standard engine control, the compressor surge characteristics, and the engine starting and windmilling characteristics. Severe circumferential turbine outlet temperature gradients which occurred at high altitude as a result of fuel nozzle coking were alleviated by the manufacturer's change in the fuel flow divider schedule and in a nozzle gasket material. Compressor air bleed is required to prevent surge of the outboard compressor in the low engine speed region. The maximum altitude at which the engine was operated without the control was about 66,000 feet at 0.8 flight Mach number and at a reduced engine speed to avoid compressor surge; with the engine control in operation, the altitude operating limit is reduced to approximately 59,000 feet. The maximum altitude at which the engine was started was about 40,000 feet.

Wallner, Lewis E.↗
NASA NTRS1947

Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. I - Pressure and Temperature Distributions

Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.

Prince, William R.↗
NASA NTRS2007

CECE: A Deep Throttling Demonstrator Cryogenic Engine for NASA's Lunar Lander

As one of the first technology development programs awarded under NASA's Vision for Space Exploration, the Pratt & Whitney Rocketdyne (PWR) Deep Throttling, Common Extensible Cryogenic Engine (CECE) program was selected by NASA in November 2004 to begin technology development and demonstration toward a deep throttling, cryogenic Lunar Lander engine for use across multiple human and robotic lunar exploration mission segments with extensibility to Mars. The CECE program leverages the maturity and previous investment of a flight-proven hydrogen/oxygen expander cycle engine, the RL10, to develop and demonstrate an unprecedented combination of reliability, safety, durability, throttlability, and restart capabilities in a high-energy, cryogenic engine. NASA Marshall Space Flight Center and NASA Glenn Research Center personnel were integral design and analysis team members throughout the requirements assessment, propellant studies and the deep throttling demonstrator elements of the program. The testbed selected for the initial deep throttling demonstration phase of this program was a minimally modified RL10 engine, allowing for maximum current production engine commonality and extensibility with minimum program cost. In just nine months from technical program start, CECE Demonstrator No. 1 engine testing in April/May 2006 at PWR's E06 test stand successfully demonstrated in excess of 10:1 throttling of the hydrogen/oxygen expander cycle engine. This test provided an early demonstration of a viable, enabling cryogenic propulsion concept with invaluable system-level technology data acquisition toward design and development risk mitigation for both the subsequent CECE Demonstrator No. 2 program and to the future Lunar Lander Design, Development, Test and Evaluation effort.

Giuliano, Victor J.↗