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Wiederien R.J.,Aerojet
47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011 | Year: 2011

Design, analysis and assembly of satellite propulsion systems typically takes 15 to 24 months from initial contract award to delivery of the system. Aerojet completed the design and analysis of a medium sized satellite propulsion system in 6 months and delivered it 12 months from the date of contract award. To meet this schedule, Aerojet and the customer formed an integrated product team with senior personnel, removed artificial boundaries to the exchange of information and focused on a simple design that used only flight qualified components. Early emphasis on trade studies with input from all relevant engineering disciplines and the customer helped to quickly identify the system architecture. Small design and manufacturing improvements that allowed more activities to be completed in parallel were implemented throughout the program and cumulatively resulted in substantial schedule reductions relative to typical program execution. The delivery schedule of purchased components was a difficult challenge and set the pace of the program. This paper explores the design and manufacturing of this satellite propulsion system to identify steps that helped facilitate a compressed delivery schedule and identify opportunities to further shorten delivery schedules on subsequent systems. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Bulman M.,Aerojet
48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2012 | Year: 2012

The effect of launch orbit for an NTR mission to Mars will be evaluated. In particular, the impact of departing from a 1000km vs. 407km circular orbit before operation of an NTR, and quantification of the mass 'cost' of departing from a 1000km. In this paper we evaluate the following options to raise the vehicle to the departure orbit: Direct ascent to 1000km by the current SLS Heavy Lift Launch Vehicle (HLV). The Assembled vehicles will be lifted to the departure orbit with a Hydrogen/Oxygen tug/kick stage. An integrated propulsion system with chemical propulsion for ascent from assembly to 1000km orbits. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Mackenzie-Helnwein P.,University of Washington | Arduino P.,University of Washington | Shin W.,University of Washington | Moore J.A.,Aerojet | Miller G.R.,University of Washington
International Journal for Numerical Methods in Engineering | Year: 2010

This paper presents an investigation of strategies for handling dissipative phase interactions in the context of multi-field material point method formulations in which each phase is assigned its own motion. Different families of phase interaction strategies using both nodal and particle-based approaches are developed, and in particular, a new smoothed volume fraction approach is presented that can handle interaction effects in a general and consistent manner while reducing anomalous effects of phase boundaries and grid crossings. The effectiveness of this approach is demonstrated via convergence studies using a fundamental model problem. © 2010 John Wiley & Sons, Ltd.

O'Brien T.F.,Aerojet
17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference 2011 | Year: 2011

A computational study of the viscous performance of a blunt leading-edged, streamlinetraced Busemann inlet with no truncation was performed. A sugar scoop style of Busemann inlet with geometric contraction ratio 7 and leading edge radius of 0.03" was sized for a nominal inviscid mass capture of 10 lbm/s at Mach 7, 0° angle of attack, and a dynamic pressure of 1000 psf. Solutions were obtained at angles of attack ranging from -5° to 10°, in increments of 5°. All angles of attack were calculated at Mach numbers ranging from 5 to 8, in increments of 1. Additional lower Mach number solutions were obtained for each angle of attack to determine operability limits for minimum running Mach number. Conservationaveraged performance results for total pressure ratio and Mach number were compiled at both the cowl lip and inlet throat planes and compared with area- and mass-averaged results. Lift, drag, and pitching moment coefficients were integrated to both the cowl lip and inlet throat planes. Performance trends were presented, showing that the inlet performance correlated well with a drag coefficient referenced to the freestream streamtube area captured by the inlet. Comparisons between the data and existing correlations for maximum contraction ratio showed good agreement when using inlet cowl plane Mach number and internal contraction ratio. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Wilson A.C.,Aerojet
61st International Astronautical Congress 2010, IAC 2010 | Year: 2010

Performance Characterization of the Advanced Materials Bipropellant Rocket (AMBR) engine was completed at Aerojet's Redmond, Washington test facility in the summer of 2009. This project was funded by the NASA In-Space Propulsion Technology (ISPT) Project Office. The primary goal of this project was to maximize the specific impulse of a pressure fed, apogee class earth storable bi-propellant engine using nitrogen tetroxide (MON-3) oxidizer and hydrazine fuel. The secondary goal of the project was to take greater advantage of the high temperature capabilities of iridium/rhenium material used for the combustion chamber and nozzle. The first round of hot fire testing of the AMBR engine occurred in October of 2008; during which, the engine demonstrated a maximum specific impulse of 333.5 seconds at a mixture ratio of 1.1 and a thrust of 151-lbf (672 N). This operating thrust level at this mixture ratio was close to the thermal stability (loss of fuel film cooling) thrust limit of the engine. A second round of testing, including random vibration, shock and extended hot fire testing, was added to the program with the goal of bringing the AMBR engine to a Technology Readiness Level (TRL) of 6. Following successful random vibration and shock testing, the AMBR engine went through a second hot fire test series in the summer of 2009 to document the performance (thrust and mixture ratio) operating margins and to demonstrate a long duration burn. During this testing, the engine demonstrated a specific impulse of 333 seconds at a mixture ratio of 1.1 and a thrust of 141-lbf (627 N) (i.e. the AMBR engine demonstrated high performance with margin). This performance also occurred during a successful 2,700 second long burn. This program also successfully demonstrated the secondary goal of hot fire operation at a 4000°F (2200°C) combustion chamber temperature. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

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