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  • 1
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    Solar Electric Propulsion (SEP) Tug Power System Considerations (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: Solar electric propulsion (SEP) technology is truly at the "intersection of commercial and military space" as well as the intersection of NASA robotic and human space missions. Building on the use of SEP for geosynchronous spacecraft station keeping, there are numerous potential commercial and military mission applications for SEP stages operating in Earth orbit. At NASA, there is a resurgence of interest in robotic SEP missions for Earth orbit raising applications, 1-AU class heliocentric missions to near Earth objects (NEOs) and SEP spacecraft technology demonstrations. Beyond these nearer term robotic missions, potential future human space flight missions to NEOs with high-power SEP stages are being considered. To enhance or enable this broad class of commercial, military and NASA missions, advancements in the power level and performance of SEP technologies are needed. This presentation will focus on design considerations for the solar photovoltaic array (PVA) and electric power system (EPS) vital to the design and operation of an SEP stage. The engineering and programmatic pros and cons of various PVA and EPS technologies and architectures will be discussed in the context of operating voltage and power levels. The impacts of PVA and EPS design options on the remaining SEP stage subsystem designs, as well as spacecraft operations, will also be discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2011-217197 , E-17882 , 2011 Space Power Workshop; 18-21 Aprl. 2011; Los Angeles, CA; United States
    Format: application/pdf
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  • 2
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    Advanced Space Transportation Concepts and Propulsion Technologies for a New Delivery Paradigm (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes Advanced Space Transportation Concepts and Propulsion Technologies for a New Delivery Paradigm. It builds on the work of the previous paper "Approach to an Affordable and Productive Space Transportation System". The scope includes both flight and ground system elements, and focuses on their compatibility and capability to achieve a technical solution that is operationally productive and also affordable. A clear and revolutionary approach, including advanced propulsion systems (advanced LOX rich booster engine concept having independent LOX and fuel cooling systems, thrust augmentation with LOX rich boost and fuel rich operation at altitude), improved vehicle concepts (autogeneous pressurization, turbo alternator for electric power during ascent, hot gases to purge system and keep moisture out), and ground delivery systems, was examined. Previous papers by the authors and other members of the Space Propulsion Synergy Team (SPST) focused on space flight system engineering methods, along with operationally efficient propulsion system concepts and technologies. This paper continues the previous work by exploring the propulsion technology aspects in more depth and how they may enable the vehicle designs from the previous paper. Subsequent papers will explore the vehicle design, the ground support system, and the operations aspects of the new delivery paradigm in greater detail.
    Keywords: Spacecraft Propulsion and Power
    Type: KSC-2013-150 , 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jun 25, 2013; San Jose, CA; United States
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  • 3
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    Summary of Results from Space Shuttle Main Engine Off-Nominal Testing (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper is a summary of Space Shuttle Main Engine (SSME) off-nominal testing that occurred during 2008 and 2009. During the last two years of planned SSME testing at Stennis Space Center, Pratt & Whitney Rocketdyne worked with their NASA MSFC customer to systematically identify, develop, assess, and implement challenging test objectives in order to expand the knowledge of one of the world s most reliable and highly tested large rocket engine. The objectives successfully investigated three main areas of interest expanding engine performance margins, demonstrating system operational capabilities, and establishing ground work for new rocket engine technology. The testing gave the Space Shuttle Program new options to safely fly out the flight manifest and provided Pratt & Whitney Rocketdyne and NASA new insight into the operational capabilities of the SSME, capabilities which can be used in assessing potential future applications of the RS-25 engine.
    Keywords: Spacecraft Propulsion and Power
    Type: M11-0276 , M11-1054 , Space 2011 AIAA Conference; Sep 27, 2011 - Sep 29, 2011; Long Beach, CA; United States
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  • 4
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    Demagnetization Tests Performed on a Linear Alternator for a Stirling Power Convertor (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The NASA Glenn Research Center (GRC) is conducting in-house research on rare-earth permanent magnets and linear alternators to assist in developing free-piston Stirling convertors for radioisotope space power systems and for developing advanced linear alternator technology. This research continues at GRC, but, with the exception of Advanced Stirling Radioisotope Generator references, the work presented in this paper was conducted in 2005. A special arc-magnet characterization fixture was designed and built to measure the M-H characteristics of the magnets used in Technology Demonstration Convertors developed under the 110-W Stirling Radioisotope Generator (SRG110) project. This fixture was used to measure these characteristics of the arc magnets and to predict alternator demagnetization temperatures in the SRG110 application. Demagnetization tests using the TDC alternator on the Alternator Test Rig were conducted for two different magnet grades: Sumitomo Neomax 44AH and 42AH. The purpose of these tests was to determine the demagnetization temperatures of the magnets for the alternator under nominal loads. Measurements made during the tests included the linear alternator terminal voltage, current, average power, magnet temperatures, and stator temperatures. The results of these tests were found to be in good agreement with predictions. Alternator demagnetization temperatures in the Advanced Stirling Convertor (ASC-developed under the Advanced Stirling Radioisotope Generator project) were predicted as well because the prediction method had been validated through the SRG110 alternator tests. These predictions led to a specification for maximum temperatures of the ASC pressure vessel.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2012-217288 , AIAA Paper 2011-5728 , E17823
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  • 5
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    Wall Pressure Unsteadiness and Side Loads in Overexpanded Rocket Nozzles (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Surveys of both the static and dynamic wall pressure signatures on the interior surface of a sub-scale, cold-flow and thrust optimized parabolic nozzle are conducted during fixed nozzle pressure ratios corresponding to FSS and RSS states. The motive is to develop a better understanding for the sources of off-axis loads during the transient start-up of overexpanded rocket nozzles. During FSS state, pressure spectra reveal frequency content resembling SWTBLI. Presumably, when the internal flow is in RSS state, separation bubbles are trapped by shocks and expansion waves; interactions between the separated flow regions and the waves produce asymmetric pressure distributions. An analysis of the azimuthal modes reveals how the breathing mode encompasses most of the resolved energy and that the side load inducing mode is coherent with the response moment measured by strain gauges mounted upstream of the nozzle on a flexible tube. Finally, the unsteady pressure is locally more energetic during RSS, albeit direct measurements of the response moments indicate higher side load activity when in FSS state. It is postulated that these discrepancies are attributed to cancellation effects between annular separation bubbles.
    Keywords: Spacecraft Propulsion and Power
    Type: M11-0498
    Format: application/pdf
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  • 6
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    Small Solar Electric Propulsion Spacecraft Concept for Near Earth Object and Inner Solar System Missions (2011)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Near Earth Objects (NEOs) and other primitive bodies are exciting targets for exploration. Not only do they provide clues to the early formation of the universe, but they also are potential resources for manned exploration as well as provide information about potential Earth hazards. As a step toward exploration outside Earth's sphere of influence, NASA is considering manned exploration to Near Earth Asteroids (NEAs), however hazard characterization of a target is important before embarking on such an undertaking. A small Solar Electric Propulsion (SEP) spacecraft would be ideally suited for this type of mission due to the high delta-V requirements, variety of potential targets and locations, and the solar energy available in the inner solar system.Spacecraft and mission trades have been performed to develop a robust spacecraft design that utilizes low cost, off-the-shelf components that could accommodate a suite of different scientific payloads for NEO characterization. Mission concepts such as multiple spacecraft each rendezvousing with different NEOs, single spacecraft rendezvousing with separate NEOs, NEO landers, as well as other inner solar system applications (Mars telecom orbiter) have been evaluated. Secondary launch opportunities using the Expendable Secondary Payload Adapter (ESPA) Grande launch adapter with unconstrained launch dates have also been examined.
    Keywords: Spacecraft Propulsion and Power
    Type: IAA Low-Cost Planetary Missions Conference; Jun 21, 2011; Laurel, MD; United States
    Format: text
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  • 7
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    Optical Mass Gauging System for Measuring Liquid Levels in a Reduced Gravity Environment (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A compact and rugged fiber-coupled liquid volume sensor designed for flight on a sounding rocket platform is presented. The sensor consists of a Mach-Zehnder interferometer capable of measuring the amount of liquid contained in a tank under any gravitational conditions, including a microgravity environment, by detecting small changes in the index of refraction of the gas contained within a sensing region. By monitoring changes in the interference fringe pattern as the system undergoes a small compression provided by a piston, the ullage volume of a tank can be directly measured allowing for a determination of the liquid volume. To demonstrate the technique, data are acquired using two tanks containing different volumes of liquid, which are representative of the levels of liquid in a tank at different time periods during a mission. The two tanks are independently exposed to the measurement apparatus, allowing for a determination of the liquid level in each. In a controlled, laboratory test of the unit, the system demonstrated a capability of measuring a liquid level in an individual tank of 10.53 mL with a 2% error. The overall random uncertainty for the flight system is higher than that one test, at +/- 1.5 mL.
    Keywords: Spacecraft Propulsion and Power
    Type: M10-0919 , 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 25, 2010 - Jul 28, 2010; Nashville, TN; United States
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  • 8
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    Details of Side Load Test Data and Analysis for a Truncated Ideal Contour Nozzle and a Parabolic Contour Nozzle (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Two cold flow subscale nozzles were tested for side load characteristics during simulated nozzle start transients. The two test article contours were a truncated ideal and a parabolic. The current paper is an extension of a 2009 AIAA JPC paper on the test results for the same two nozzle test articles. The side load moments were measured with the strain tube approach in MSFC s Nozzle Test Facility. The processing techniques implemented to convert the strain gage signals into side load moment data are explained. Nozzle wall pressure profiles for separated nozzle flow at many NPRs are presented and discussed in detail. The effect of the test cell diffuser inlet on the parabolic nozzle s wall pressure profiles for separated flow is shown. The maximum measured side load moments for the two contours are compared. The truncated ideal contour s peak side load moment was 45% of that of the parabolic contour. The calculated side load moments, via mean-plus-three-standard-deviations at each nozzle pressure ratio, reproduced the characteristics and absolute values of measured maximums for both contours. The effect of facility vibration on the measured side load moments is quantified and the effect on uncertainty is calculated. The nozzle contour designs are discussed and the impact of a minor fabrication flaw in the nozzle contours is explained.
    Keywords: Spacecraft Propulsion and Power
    Type: M10-0862 , AIAA Joint Propulsion Conference; Jul 26, 2010 - Jul 28, 2010; Nashville, TN; United States
    Format: application/pdf
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  • 9
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    Performance Testing of a Prototypic Annular Linear Induction Pump for Fission Surface Power (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Results of performance testing of an annular linear induction pump are presented. The pump electromagnetically pumps liquid metal (NaK) through a circuit specially designed to allow for quantification of the performance. Testing was conducted over a range of conditions, including frequencies of 33, 36, 39, and 60 Hz, liquid metal temperatures from 25 to 525 C, and input voltages from 5 to 120 V. Pump performance spanned a range of flow rates from roughly 0.16 to 5.7 L/s (2.5 to 90 gpm), and pressure head 〈1 to 90 kPa (〈0.145 to 13 psi). The maximum efficiency measured during testing was slightly greater than 6%. The efficiency was fairly insensitive to input frequency from 33 to 39 Hz, and was markedly lower at 60 Hz. In addition, the efficiency decreased as the NaK temperature was raised. While the pump was powered, the fluid responded immediately to changes in the input power level, but when power was removed altogether, there was a brief slow-down period before the fluid would come to rest. The performance of the pump operating on a variable frequency drive providing 60 Hz power compared favorably with the same pump operating on 60 Hz power drawn directly from the electrical grid.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TP-2010-216430 , M-1279
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  • 10
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    Compact High Current Rare-Earth Emitter Hollow Cathode for Hall Effect Thrusters (2012)
    In:  CASI
    Details
    Publication Date: 2019-08-24
    Description: An apparatus and method for achieving an efficient central cathode in a Hall effect thruster is disclosed. A hollow insert disposed inside the end of a hollow conductive cathode comprises a rare-earth element and energized to emit electrons from an inner surface. The cathode employs an end opening having an area at least as large as the internal cross sectional area of the rare earth insert to enhance throughput from the cathode end. In addition, the cathode employs a high aspect ratio geometry based on the cathode length to width which mitigates heat transfer from the end. A gas flow through the cathode and insert may be impinged by the emitted electrons to yield a plasma. One or more optional auxiliary gas feeds may also be employed between the cathode and keeper wall and external to the keeper near the outlet.
    Keywords: Spacecraft Propulsion and Power
    Format: application/pdf
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