SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 13 - JULY 5, 2006
20 SPACECRAFT PROPULSION AND POWER
Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxiliary power sources.
For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 15 Launch Vehicles and Launch operations, and 44 Energy Production and Conversion.
20060017028 NASA Marshall Space Flight Center, Huntsville, AL, USA
Marshall Space Flight Center Propulsion Systems Department (PSD) KM Initiative
Caraccioli, Paul; Varnadoe, Tom; McCarter, Mike; [2006]; 2 pp.; In English; Managing Knowledge for Successful Mission Operations, 2-3 Mar. 2006, Houston, TX, USA Contract(s)/Grant(s): NNM05AA44G; No Copyright; Avail.: Other Sources; Abstract Only
NASA Marshall Space Flight Center s Propulsion Systems Department (PSD) is four months into a fifteen month Knowledge Management (KM) initiative to support enhanced engineering decision making and analyses, faster resolution of anomalies (near-term) and effective, efficient knowledge infused engineering processes, reduced knowledge attrition, and reduced anomaly occurrences (long-term). The near-term objective of this initiative is developing a KM Pilot project, within the context of a 3-5 year KM strategy, to introduce and evaluate the use of KM within PSD. An internal NASA/MSFC PSD KM team was established early in project formulation to maintain a practitioner, user-centric focus throughout the conceptual development, planning and deployment of KM technologies and capabilities with in the PSD. The PSD internal team is supported by the University of Alabama's Aging Infrastructure Systems Center Of Excellence (AISCE), Intergraph Corporation, and The Knowledge Institute.
The principle product of the initial four month effort has been strategic planning of PSD KM implementation by first determining the 'as is' state of KM capabilities and developing, planning and documenting the roadmap to achieve the desired 'to be' state. Activities undertaken to support the planning phase have included data gathering; cultural surveys, group work-sessions, interviews, documentation review, and independent research. Assessments and analyses have been performed including industry benchmarking, related local and Agency initiatives, specific tools and techniques used and strategies for leveraging existing resources, people and technology to achieve common KM goals. Key findings captured in the PSD KM Strategic Plan include the system vision, purpose, stakeholders, prioritized strategic objectives mapped to the top ten practitioner needs and analysis of current resource usage.
Opportunities identified from research, analyses, cultural/KM surveys and practitioner interviews include: executive and senior management sponsorship, KM awareness, promotion and training, cultural change management, process improvement, leveraging existing resources and new innovative technologies to align with other NASA KM initiatives (convergence: the big picture). To enable results based incremental implementation and future growth of the KM initiative, key performance measures have been identified including stakeholder value, system utility, learning and growth (knowledge capture, sharing, reduced anomaly recurrence), cultural change, process improvement and return-on-investment. The next steps for the initial implementation spiral (focused on SSME Turbomachinery) have been identified, largely based on the organization and compilation of summary level engineering process models, data capture matrices, functional models and conceptual-level systems architecture.
Key elements includedetailed KM requirements definition, KM technology architecture assessment, evaluation and selection, deployable KM Pilot design, development, implementation and evaluation, and justifying full implementation (estimated Return-on-Investment). Features identified for the notional system architecture include the knowledge presentation layer (and its components), knowledge network layer (and its components), knowledge storage layer (and its components), User Interface and capabilities. This paper provides a snapshot of the progress to date, the near term planning for deploying the KM pilot project and a forward look at results based growth of KM capabilities with-in the MSFC PSD. Derived from text
Knowledge Representation; Management Planning; Propulsion System Configurations; Decision Making; Expert Systems; Deployment; Propulsion System Performance
20060017559 Photonic Associates, Santa Fe, NM USA
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3ks Specific Impulse with a ns-pulse Laser Microthruster
Phipps, Claude R; Luke, James R; Helgeson, Wesley D; Aug 23, 2005; 11 pp.; In English Contract(s)/Grant(s): Proj-BMSB Report No.(s): AD-A444975; No Copyright; Avail.: CASI: A03, Hardcopy
Photonics developed a prototype device which demonstrated the feasibility of using ns-duration laser pulses in a laser microthruster. Using several metallic targets driven by a 'microchip' laser, thrust, specific impulse Isp, specific ablation energy Q* and ablation efficiency were measured. Specific impulse was adjusted by varying laser intensity on target. In this way, researchers were able to vary specific impulse from 200s to 3,800s on gold, with corresponding momentum coupling coefficient Cm varying from 70 to 7 micro-N/optical watt and ablation efficiency near 100% at a 1 MJK/sq m optimum pulse fluence. Simulations of the laser-target interaction are discussed, which agreed well with the results obtained on metallic fuel systems. A Concepts Research, Inc., microchip laser was used with 170mW average optical power, 8kHz repetition rate and 20 micro-J pulse energy for many of the measurements. Thrust was in the 100nN - 1 micro-N range for all the work, requiring development of an extremely sensitive, low-noise thrust stand, which is discussed in a companion paper. The design of realistic metallic fuel delivery systems is discussed.
Also reported are time-of-flight measurements on ejected metal ions, which gave velocities up to 80km/s. Near the optimum fluence, good agreement was found between Isp deduced from the CmQ* product and from ion velocity v(sub-i). Strong divergence was observed between these two parameters at higher intensity. Relative to the ms-duration thrusters which have been demonstrated in the past, this change offers the use of any target material, the use of more-efficient reflection-mode target illumination, and specific impulse adjustable to match the mission. DTIC
Electric Propulsion; Microrocket Engines; Pulsed Lasers; Spacecraft Propulsion; Specific Impulse
20060017585 Northrop Grumman Space and Mission Systems Corp., Redondo Beach, CA USA
Experimental and Numerical Analysis of Transpiration Cooling of a Rocket Engine Using Lamilloy (Trademark) Plates (POSTPRINT)
Auyeung, T P; Cohn, R; Coy, E; Danczyk, S A; Papesh, C; Sweeney, P; Dec 2005; 13 pp.; In English Contract(s)/Grant(s): NRO000-02-05-0516; Proj-5026 Report No.(s): AD-A445014; No Copyright; Avail.: CASI: A03, Hardcopy
Transpiration cooling of rocket engine thrust chamber walls has the potential for improving the performance of liquid rocket engines by reducing the required cooling flow. For the TR107 hydrocarbon booster engine, it is estimated that replacing the proposed chamber wall film cooling with transpiration cooling could result in an engine Isp increase of 2% to 3%, which provides the potential to increase the maximum booster delivery capability by 4000-5000 lbm. In the past, various transpiration cooling wall materials have been investigated with varying degrees of success. In this investigation we examined the use of Lamilloy * as the transpirant wall. Lamilloy is a cooling system with a 30+ year history in jet engine applications developed by Rolls-Royce LibertyWorks. This work represents the first evaluation of this technology for rocket engine applications in a representative hot-fire environment. Tests were performed in a sub-scale hot-fire chamber that produced heat fluxes of 4-15 Btu/in.2/s at chamber pressures of 370-670 psi. Gaseous oxygen/RP-1 propellants were fired at mixture ratios ranging from 1.2 to 1.8. Three off-the-shelf Lamilloy designs were tested with three different transpirants: gaseous nitrogen, water, and RP-1. Testing demonstrated that the hot-gas wall temperature decreased rapidly in the downstream direction due to the cumulative effect of the injected coolant. The tested Lamilloy specimens demonstrated the potential to be an effective material for use as a transpiration cooled wall. However, the specimens tested were designed for low pressure drop in gas turbine combustor applications thus did not provide the high pressure drop needed for rocket propulsion applications. The extensive database generated during this testing can be used to guide future Lamilloy transpiration cooling designs for specific rocket thrust chamber applicat DTIC
Booster Rocket Engines; Film Cooling; Heat Resistant Alloys; Liquid Propellant Rocket Engines; Numerical Analysis; Rocket Engines; Sweat Cooling; Transpiration
20060017599 Engineering Research and Consulting, Inc., Edwards AFB, CA USA
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Experiments on a Coaxial Injector Under an Externally-Forced Transverse Acoustic Field (POSTPRINT)
Davis, D W; Chehroudi, B; Talley, D G; Jun 20, 2005; 21 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A445032; AFRL-PR-ED-TP-2005-410; No Copyright; Avail.: CASI: A03, Hardcopy
In order to gain a better understanding of some of the underlying physics associated with the interaction of high-amplitude acoustic waves and a coaxial-jet type injector similar to those used in cryogenic liquid rockets, a non-reacting-flow experimental investigation was conducted under sub-, near-, and supercritical chamber pressures, with and without acoustical excitation. Past research works on this subject have shown both the relevance and importance of geometrical changes in an injector's exit-area and its nearby physical and fluid mechanical processes. On this basis, special attention is paid in collecting spatially-resolved mean temperatures and documenting the aforementioned interactions at the exit of this injector. Short-duration and high-speed digital cameras provided information on the dynamic behavior of this jet. Mean and root mean square (RMS) values of the coaxial-jet dark-core length fluctuations were measured from the acquired images via a computer-automated method. It is seen that as the outer-to-inner jet velocity ratio increases, the RMS of the dark-core length fluctuations decreases.
It is hypothesized that a connection to rocket instability can be obtained from the data analyzed thus far by way of the magnitude of the RMS values of the dark-core length fluctuations. It is possible that decreases in the fluctuation levels, which were shown here to occur at higher velocity ratios, could weaken a key feedback mechanism for the self-excitation process that is believed to drive the combustion instability in rocket engines. This could offer a possible explanation of the combustion stability improvements experienced in engines under higher outer-to-inner jet velocity ratios. Finally, there also appears to be a good correlation between the dark-core length and the outer-to-inner jet momentum flux ratio, but the form of this dependence was found to be different at subcritical pressures than the rest of the chamber conditions. DTIC
Acoustics; Injectors; Liquid Propellant Rocket Engines; Sound Fields
20060017638 Army Armament Research, Development and Engineering Center, Picatinny Arsenal, NJ USA
M549A1 Projectile Delay Assembly Predictive Engineering Analysis in Support of the Ammunition Stockpile Reliability Program
Bixon, Eric R; Cassiello, Michael J; McEwan, Jr, John D; Mar 2006; 30 pp.; In English Report No.(s): AD-A445121; ARQES-TR-06001; No Copyright; Avail.: CASI: A03, Hardcopy
Experiments were conducted on M549A1 High Explosive Rocket Assisted projectiles to evaluate the effects of thermal degradation and moisture on the delay assembly, and on the potential for premature rocket ignitions. Accelerated aging experiments for up to 25 days at 75 degrees F did not induce premature rocket motor ignitions. Higher temperatures and conditioning times were recommended in order to cause this type of failure mechanism to become active. Delay assemblies required very high temperatures (375 degrees F and 475 degrees F) to cause failure due to temperature alone. When humiditywas introduced into the test matrix, the delays failed at much lower temperatures (1 55 degrees F to 230 degrees F). Because of the high susceptibility to failure of the delay assembly due to humidity exposure, the recommendation was made not to fire any M549A1 projectiles which do not have the Rocket Off Cap present on the projectile. DTIC
Ammunition; Predictions; Projectiles; Reliability; Rocket Engines; Stockpiling
20060017673 Air Force Inst. of Tech., Wright-Patterson AFB, OH USA
A Prediction Code for the Thrust Performance of Two-Dimensional, Non-Axisynnetric, Converging Diverging Nozzles
Geatz, Angela M; Sep 2005; 119 pp.; In English; Original contains color illustrations Report No.(s): AD-A445183; AFIT/GAE/ENY/06-03; No Copyright; Avail.: CASI: A06, Hardcopy
The objective of this research is to develop a prediction code for the Air Force Research Laboratory Propulsion Directorate that can accurately determine the gross thrust coefficient for a user defined nonaxisymmetric two-dimensional converging diverging nozzle. The code includes the effects of friction, angularity, and expansion losses on nozzle efficiency. To demonstrate the prediction method, the generated computational results were compared to experimental data, as well as computational results from other existing nozzle performance codes, for a number of different nozzle geometries. The nozzle internal performance prediction code showed excellent agreement with experimental data in predicting the gross thrust performance for all nozzle geometries considered. It was shown, however, that when the experimental data showed evidence of flow separation, a flow phenomenon this code is unable to predict, the code results underpredicted the experimental by up to 10%. DTIC
Convergent-Divergent Nozzles; Thrust
20060018509 Universidad Politecnica de Madrid, Madrid, Spain
Numerical Modeling of the Hall Thruster Discharge
Parra, Felix I; Ahedo, Eduardo; Martinez-Sanchez, Manuel; Fife, John Michael; Escobar, D; Rus, Juan; Lapuerta, V; Molina, A; Apr 1, 2005; 119 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): FA8655-04-1-3003 Report No.(s): AD-A445352; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA445352; Avail.: CASI: A06, Hardcopy
This collection of seven previously published papers performed under Grant No. FA8655-04-1-3003 provide the background for the development of a new version of the HPHall hybrid code (HPHallv.2) for the numerical modeling of Hall Thruster discharge and new insights on discharge physics obtained during the development. DTIC
Hall Thrusters; Mathematical Models
Source: NASA
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