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NASA NTRS1995

NASA/DOD Aerospace Knowledge Diffusion Research Project. Paper 51: Workplace communications skills and the value of communications and information-use skills instruction: Engineering students' perspectives

Studies indicate that communications and information-related activities take up a substantial portion of an engineer's work week; therefore, effective communications and information use skills are one of the key engineering competencies that recent graduates of engineering programs are expected to possess. Feedback from industry rates communications and information use skills of entry-level engineers low. Missing from current discussions of communications and information use skills and competencies for engineering students is a clear explanation from the professional engineering community about what constitutes 'acceptable and desirable communications and information norms' within that community. To gather adequate and generalizable data about communications and information skills instruction and to provide a student perspective on the communications skills of engineers, we undertook a national study of aerospace engineering students in March 1993. The study included questions about the importance of certain communications and information skills to professional success, the instruction students had received in these skills, and perceived helpfulness of the instruction. Selected results from the study study are reported in this paper.

Pinelli, Thomas E.↗
NASA NTRS1995

NASA Systems Engineering Handbook

This handbook brings the fundamental concepts and techniques of systems engineering to NASA personnel in a way that recognizes the nature of NASA systems and environment. It is intended to accompany formal NASA training courses on systems engineering and project management when appropriate, and is designed to be a top-level overview. The concepts were drawn from NASA field center handbooks, NMI's/NHB's, the work of the NASA-wide Systems Engineering Working Group and the Systems Engineering Process Improvement Task team, several non-NASA textbooks and guides, and material from independent systems engineering courses taught to NASA personnel. Five core chapters cover systems engineering fundamentals, the NASA Project Cycle, management issues in systems engineering, systems analysis and modeling, and specialty engineering integration. It is not intended as a directive. Superseded by: NASA/SP-2007-6105 Rev 1 (20080008301).

Shishko, Robert↗
NASA NTRS1998

Teaching Engineering Students Team Work

The purpose of this manual is to provide professor's in engineering classes which the background necessary to use student team projects effectively. This manual describes some of the characteristics of student teams and how to use them in class. It provides a set of class activities and films which can be used to introduce and support student teams. Finally, a set of teaching modules used in freshmen, sophomore, and senior aeronautical engineering classes are presented. This manual was developed as part of a NASA sponsored project to improve the undergraduate education of aeronautical engineers. The project has helped to purchase a set of team work films which can be checked out from Cal Poly's Learning Resources Center in the Kennedy Library. Research for this project has included literature reviews on team work and cooperative learning; interviews, observations, and surveys of Cal Poly students from Industrial and Manufacturing Engineering, Aeronautical Engineering and Psychology; participation in the Aeronautical Engineering senior design lab; and interviews with engineering faculty. In addition to this faculty manual, there is a student team work manual which has been designed to help engineering students work better in teams.

Levi, Daniel↗
NASA NTRS2002

Weibull-Based Design Methodology for Rotating Aircraft Engine Structures

The NASA Energy Efficient Engine (E(sup 3)-Engine) is used as the basis of a Weibull-based life and reliability analysis. Each component's life and thus the engine's life is defined by high-cycle fatigue (HCF) or low-cycle fatigue (LCF). Knowing the cumulative life distribution of each of the components making up the engine as represented by a Weibull slope is a prerequisite to predicting the life and reliability of the entire engine. As the engine Weibull slope increases, the predicted lives decrease. The predicted engine lives L(sub 5) (95 % probability of survival) of approximately 17,000 and 32,000 hr do correlate with current engine maintenance practices without and with refurbishment. respectively. The individual high pressure turbine (HPT) blade lives necessary to obtain a blade system life L(sub 0.1) (99.9 % probability of survival) of 9000 hr for Weibull slopes of 3, 6 and 9, are 47,391 and 20,652 and 15,658 hr, respectively. For a design life of the HPT disks having probable points of failure equal to or greater than 36,000 hr at a probability of survival of 99.9 %, the predicted disk system life L(sub 0.1) can vary from 9,408 to 24,911 hr.

Zaretsky, Erwin↗
NASA NTRS2003

Sensor Based Engine Life Calculation: A Probabilistic Perspective

It is generally known that an engine component will accumulate damage (life usage) during its lifetime of use in a harsh operating environment. The commonly used cycle count for engine component usage monitoring has an inherent range of uncertainty which can be overly costly or potentially less safe from an operational standpoint. With the advance of computer technology, engine operation modeling, and the understanding of damage accumulation physics, it is possible (and desirable) to use the available sensor information to make a more accurate assessment of engine component usage. This paper describes a probabilistic approach to quantify the effects of engine operating parameter uncertainties on the thermomechanical fatigue (TMF) life of a selected engine part. A closed-loop engine simulation with a TMF life model is used to calculate the life consumption of different mission cycles. A Monte Carlo simulation approach is used to generate the statistical life usage profile for different operating assumptions. The probabilities of failure of different operating conditions are compared to illustrate the importance of the engine component life calculation using sensor information. The results of this study clearly show that a sensor-based life cycle calculation can greatly reduce the risk of component failure as well as extend on-wing component life by avoiding unnecessary maintenance actions.

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

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 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.↗
NASA NTRS2009

An Object-Oriented Computer Code for Aircraft Engine Weight Estimation

Reliable engine-weight estimation at the conceptual design stage is critical to the development of new aircraft engines. It helps to identify the best engine concept amongst several candidates. At NASA Glenn Research Center (GRC), the Weight Analysis of Turbine Engines (WATE) computer code, originally developed by Boeing Aircraft, has been used to estimate the engine weight of various conceptual engine designs. The code, written in FORTRAN, was originally developed for NASA in 1979. Since then, substantial improvements have been made to the code to improve the weight calculations for most of the engine components. Most recently, to improve the maintainability and extensibility of WATE, the FORTRAN code has been converted into an object-oriented version. The conversion was done within the NASA's NPSS (Numerical Propulsion System Simulation) framework. This enables WATE to interact seamlessly with the thermodynamic cycle model which provides component flow data such as airflows, temperatures, and pressures, etc., that are required for sizing the components and weight calculations. The tighter integration between the NPSS and WATE would greatly enhance system-level analysis and optimization capabilities. It also would facilitate the enhancement of the WATE code for next-generation aircraft and space propulsion systems. In this paper, the architecture of the object-oriented WATE code (or WATE++) is described. Both the FORTRAN and object-oriented versions of the code are employed to compute the dimensions and weight of a 300-passenger aircraft engine (GE90 class). Both versions of the code produce essentially identical results as should be the case.

Tong, Michael T.↗
NASA NTRS2010

Unleashing Lessons: Sharing Stories About the Fine Art of Systems Engineering

NASA leaders have a responsibility to share their unique oral histories with junior-level employees on whom NASA's future depends. This presentation will give a few examples of how the imaginative, flexible art of systems engineering is as necessary to mission success as is the rigorous, disciplined side of engineering. Engineering space systems involves many disciplines propulsion, loads, dynamics, and so forth that are based on the foundations of scientific principles and methodology and the application of the laws of physics. The term rocket scientist is an apt term, considering that the underlying chemical properties of propellants and the subatomic properties of materials must be understood to harness the powerful energy necessary to escape Earth's gravity in machines that can withstand the stresses and forces to which they are subjected, not to mention the harsh space environments in which they must work. This is a simplistic, yet illustrative, explanation of the scientific side of the engineer s challenge. Bringing together these individual parts into a solid system goes beyond the science of engineering to employ the art of systems engineering. Systems engineers are known for their ability to integrate various solutions to meet or exceed challenging requirements. As the old adage goes, measure twice and cut once. The act of measuring is balancing rigid, inflexible requirements with creative compromises to attain the optimum solution to the challenge of space flight. Then, we cut out those answers that are too risky, expensive, dangerous, and so forth. The process of sharing stories about the little-discussed art of engineering, also known as the art of compromise, will equip the workforce to subjectively judge the best right answer from among the many presented, while objectively integrating the various piece parts into a unified whole.

Singer, Christopher E.↗