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NASA NTRS2008

High Velocity Linear Induction Launcher with Exit-Edge Compensation for Testing of Aerospace Components

Advances in ultra high speed linear induction electromagnetic launchers over the past decade have focused on magnetic compensation of the exit and entry-edge transient flux wave to produce efficient and compact linear electric machinery. The paper discusses two approaches to edge compensation in long-stator induction catapults with typical end speeds of 150 to 1,500 m/s. In classical linear induction machines, the exit-edge effect is manifest as two auxiliary traveling waves that produce a magnetic drag on the projectile and a loss of magnetic flux over the main surface of the machine. In the new design for the Stator Compensated Induction Machine (SCIM) high velocity launcher, the exit-edge effect is nulled by a dual wavelength machine or alternately the airgap flux is peaked at a location prior to the exit edge. A four (4) stage LIM catapult is presently being constructed for 180 m/s end speed operation using double-sided longitudinal flux machines. Advanced exit and entry edge compensation is being used to maximize system efficiency, and minimize stray heating of the reaction armature. Each stage will output approximately 60 kN of force and produce over 500 G s of acceleration on the armature. The advantage of this design is there is no ablation to the projectile and no sliding contacts, allowing repeated firing of the launcher without maintenance of any sort. The paper shows results of a parametric study for 500 m/s and 1,500 m/s linear induction launchers incorporating two of the latest compensation techniques for an air-core stator primary and an iron-core primary winding. Typical thrust densities for these machines are in the range of 150 kN/sq.m. to 225 kN/sq.m. and these compete favorably with permanent magnet linear synchronous machines. The operational advantages of the high speed SCIM launcher are shown by eliminating the need for pole-angle position sensors as would be required by synchronous systems. The stator power factor is also improved.

Kuznetsov, Stephen↗

NASA NTRS1983

Auxiliary propulsion requirements for large space systems

An insight into auxiliary propulsion systems (APS) requirements for large space systems (LSS) launchable by a single shuttle is presented. In an effort to scope the APS requirements for LSS, a set of generic LSSs were defined. For each generic LSS class a specific structural configuration, representative of that most likely to serve the needs of the 1980's and 1990's was defined. The environmental disturbance forces and torques which would be acting on each specific structural configuration in LEO and GEO orbits were then determined. Auxiliary propulsion requirements were determined as a function of: generic class specific configuration, size and openness of structure, orbit, angle of orientation, correction frequency, duty cycle, number and location of thrusters and direction of thrusters and APS/LSS interactions. The results of this analysis were used to define the APS characteristics of: (1) number and distribution of thrusters, (2) thruster modulation, (3) thrust level, (4) mission energy requirements, (5) total APS mass component breakdown, and (6) state of the art adequacy/deficiency.

Maloy, J. E.↗

NASA NTRS1983

Auxiliary propulsion requirements for large space systems

Maloy, J. E.↗

NASA NTRS1983

Study of auxiliary propulsion requirements for large space systems. Volume 1: Executive summary

Smith, W. W.↗

NASA NTRS2003

Glacier Ice Mass Fluctuations and Fault Instability in Tectonically Active Southern Alaska

Across southern Alaska the northwest directed subduction of the Pacific plate is accompanied by accretion of the Yakutat terrane to continental Alaska. This has led to high tectonic strain rates and dramatic topographic relief of more than 5000 meters within 15 km of the Gulf of Alaska coast. The glaciers of this area are extensive and include large glaciers undergoing wastage (glacier retreat and thinning) and surges. The large glacier ice mass changes perturb the tectonic rate of deformation at a variety of temporal and spatial scales. We estimated surface displacements and stresses associated with ice mass fluctuations and tectonic loading by examining GPS geodetic observations and numerical model predictions. Although the glacial fluctuations perturb the tectonic stress field, especially at shallow depths, the largest contribution to ongoing crustal deformation is horizontal tectonic strain due to plate convergence. Tectonic forces are thus the primary force responsible for major earthquakes. However, for geodetic sites located < 10-20 km from major ice mass fluctuations, the changes of the solid Earth due to ice loading and unloading are an important aspect of interpreting geodetic results. The ice changes associated with Bering Glacier s most recent surge cycle are large enough to cause discernible surface displacements. Additionally, ice mass fluctuations associated with the surge cycle can modify the short-term seismicity rates in a local region. For the thrust faulting environment of the study region a large decrease in ice load may cause an increase in seismic rate in a region close to failure whereas ice loading may inhibit thrust faulting.

SauberRosenberg, Jeanne M.↗

NASA NTRS2018

Pulsed Fission-Fusion (PuFF)

In September 2013 the NASA Innovative Advanced Concept (NIAC) organization awarded a phase I contract to the PuFF team. Our phase 1 proposal discussed a pulsed fission-fusion propulsion system that injected gaseous deuterium (D) and tritium (T) as a mixture in a column, surrounded concentrically by gaseous uranium fluoride (UF6) and then an outer shell of liquid lithium. A high power current would flow down the liquid lithium and the resulting Lorentz force would compress the column by roughly a factor of 10. The compressed column would reach criticality and a combination of fission and fusion reactions would occur. The fission reactions would further energize the fusion center, and the fusion reactions would generate neutrons that promote more complete burnup of the fission fuel. The lithium liner provides some help as a neutron reflector but also acts as a propulsive medium, being converted to plasma which is then expanded against a magnetic nozzle for thrust. The expansion of the (primarily) lithium plasma against the nozzle's magnetic field inducts a current that is used to charge the system for the next pulse. Our concept also included secondary injection of a Field Reversed Configuration (FRC) plasmoid that would provide a secondary compression direction, axially against the column, and push the column away from the injection manifold, increasing the manifold's survivability.Our phase 1 proposal included modeling the above process first under steady state assumptions and second under a time variant integration. We proposed including these results into a Mars concept vehicle and finally proposing promising conditions to be evaluated experimentally in Phase II. In phase I we quickly realized that we needed to modify our approach. Our steady state work was completed as proposed, and the results indicated that one, a two stage compression system was not needed and two, that we wanted to move away from UF6. The steady state model shows much more margin than expected, to the point that we may well reach breakeven with the Charger â€“ 1 facility, a 572 kJ Marx bank currently under refurbishment at UAH. Additionally we found that using gaseous D-T and UF6, provided a relatively simple prospect of using a pulsed injector, made reaching criticality more difficult. The introduction of large amounts of fluorine meant a radiative sink, sapping power from the fusion plasma and was harder to handle. Therefore we moved to a solid uranium target that held D-T under pressure. In so doing we could move our target closer to criticality and remove any material that did not sustain the reaction.

Fusion↗

NASA NTRS1995

Low Noise Research Fan Stage Design

This report describes the design of a Low Noise ADP Research Fan stage. The fan is a variable pitch design which is designed at the cruise pitch condition. Relative to the cruise setting, the blade is closed at takeoff and opened for reverse thrust operation. The fan stage is a split flow design with fan exit guide vanes and core stators. This fan stage design was combined with a nacelle and engine core duct to form a powered fan/nacelle, subscale model. This model is intended for use in aerodynamic performance, acoustic and structural testing in a wind tunnel. The model has a 22-inch outer fan diameter and a hub-to-top ratio of 0.426 which permits the use of existing NASA fan and cowl force balance designs and rig drive system. The design parameters were selected to permit valid acoustic and aerodynamic comparisons with the PW 17-inch rig previously tested under NASA contract. The fan stage design is described in detail. The results of the design axisymmetric analysis at aerodynamic design condition are included. The structural analysis of the fan rotor and attachment is described including the material selections and stress analysis. The blade and attachment are predicted to have adequate low cycle fatigue life, and an acceptable operating range without resonant stress or flutter. The stage was acoustically designed with airfoil counts in the fan exit guide vane and core stator to minimize noise. A fan-FEGV tone analysis developed separately under NASA contract was used to determine these airfoil counts. The fan stage design was matched to a nacelle design to form a fan/nacelle model for wind tunnel testing. The nacelle design was developed under a separate NASA contract. The nacelle was designed with an axisymmetric inlet, cowl and nozzle for convenience in testing and fabrication. Aerodynamic analysis of the nacelle confirmed the required performance at various aircraft operating conditions.

Hobbs, David E.↗

NASA NTRS1997

Advanced Low-Noise Research Fan Stage Design

This report describes the design of the Advanced Low-Noise Research Fan stage. The fan is a variable pitch design, which is designed at the cruise pitch condition. Relative to the cruise setting, the blade is closed at takeoff and opened for reverse thrust operation. The fan stage is a split flow design with fan exit guide vanes (FEGVs) and core stators. The fan stage design is combined with a nacelle and engine core duct to form a powered fan/nacelle subscale model. This model is intended for use in combined aerodynamic, acoustic, and structural testing in a wind tunnel. The fan has an outer diameter of 22 in. and a hub-to-tip of 0.426 in., which allows the use of existing NASA fan and cowl force balance and rig drive systems. The design parameters were selected to permit valid acoustic and aerodynamic comparisons with the Pratt & Whitney (P&W) 17- and 22-in. rigs previously tested under NASA contract. The fan stage design is described in detail. The results of the design axisymmetric and Navier-Stokes aerodynamic analysis are presented at the critical design conditions. The structural analysis of the fan rotor and attachment is included. The blade and attachment are predicted to have adequate low-cycle fatigue life and an acceptable operating range without resonant stress or flutter. The stage was acoustically designed with airfoil counts in the FEGV and core stator to minimize noise. A fan/FEGV tone analysis developed separately under NASA contract was used to determine the optimum airfoil counts. The fan stage was matched to the existing nacelle, designed under the previous P&W low-noise contract, to form a fan/nacelle model for wind tunnel testing. It is an axisymmetric nacelle for convenience in testing and analysis. Previous testing confirmed that the nacelle performed as required at various aircraft operating conditions.

Neubert, Robert↗

NASA NTRS1981

JT9D performance deterioration results from a simulated aerodynamic load test

The results of testing to identify the effects of simulated aerodynamic flight loads on JT9D engine performance are presented. The test results were also used to refine previous analytical studies on the impact of aerodynamic flight loads on performance losses. To accomplish these objectives, a JT9D-7AH engine was assembled with average production clearances and new seals as well as extensive instrumentation to monitor engine performance, case temperatures, and blade tip clearance changes. A special loading device was designed and constructed to permit application of known moments and shear forces to the engine by the use of cables placed around the flight inlet. The test was conducted in the Pratt & Whitney Aircraft X-Ray Test Facility to permit the use of X-ray techniques in conjunction with laser blade tip proximity probes to monitor important engine clearance changes. Upon completion of the test program, the test engine was disassembled, and the condition of gas path parts and final clearances were documented. The test results indicate that the engine lost 1.1 percent in thrust specific fuel consumption (TSFC), as measured under sea level static conditions, due to increased operating clearances caused by simulated flight loads. This compares with 0.9 percent predicted by the analytical model and previous study efforts.

Stakolich, E. G.↗

NASA NTRS1988

A cyclic ground test of an ion auxiliary propulsion system: Description and operational considerations

The Ion Auxiliary Propulsion System (IAPS) experiment is designed for launch on an Air Force Space Test Program satellite (NASA-TM-78859; AIAA Paper No. 78-647). The primary objective of the experiment is to flight qualify the 8 cm mercury ion thruster system for stationkeeping applications. Secondary objectives are measuring the interactions between operating ion thruster systems and host spacecraft, and confirming the design performance of the thruster systems. Two complete 8 cm mercury ion thruster subsystems will be flown. One of these will be operated for 2557 on and off cycles and 7057 hours at full thrust. Tests are currently under way in support of the IAPS flight experiment. In this test an IAPS thruster is being operated through a series of startup/run/shut-down cycles which simulate thruster operation during the planned flight experiment. A test facility description and operational considerations of this testing using an engineering model 8 cm thruster (S/N 905) is the subject of this paper. Final results will be published at a later date when the ground test has been concluded.

Ling, Jerri S.↗

NASA NTRS1987

Large-scale Advanced Prop-fan (LAP) static rotor test report

Discussed is Static Rotor Testing of the SR-7L Large Scale Advanced Prop-Fan (LAP). The LAP is an advanced 9 foot diameter, 8 bladed propeller designed and built by Hamilton Standard under contract to the NASA Lewis Research Center. The Prop-Fan employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Static Testing was conducted on a 10,000 HP whirl rig at Wright Patterson Air Force Base. The test objectives were to investigate the Prop-Fan static aerodynamic and structural dynamic performance, determine the blade steady state stressers and deflections and to measure steady and unsteady pressures on the SR-7L blade surface. The measured performance of the LAP correlated well with analytical predictions at blade pitch angles below 30 deg. A stall buffet phenomenon was observed at blade pitch angles above 30 deg. This phenomenon manifested itself by elevated blade vibratory stress levels and lower than expected thrust produced and power absorbed by the Prop-Fan for a given speed and blade angle.

Degeorge, Charles L.↗

NASA NTRS1981

JT9D performance deterioration results from a simulated aerodynamic load test

This paper presents the results of testing to identify the effects of simulated aerodynamic flight loads on JT9D engine performance. The test results were also used to refine previous analytical studies on the impact of aerodynamic flight loads on performance losses. To accomplish these objectives, a JT9D-7AH engine was assembled with average production clearances and new seals as well as extensive instrumentation to monitor engine performance, case temperatures, and blade tip clearance changes. A special loading device was designed and constructed to permit application of known moments and shear forces to the engine by the use of cables placed around the flight inlet. The test was conducted in the Pratt and Whitney Aircraft X-Ray Test Facility to permit the use of X-ray techniques in conjunction with laser blade tip proximity probes to monitor important engine clearance changes. Upon completion of the test program, the test engine was disassembled, and the condition of gas path parts and final clearances were documented. The test results indicate that the engine lost 1.1 percent in thrust specific fuel consumption (TSFC), as measured under sea level static conditions, due to increased operating clearances caused by simulated flight loads. This compares with 0.9 percent predicted by the analytical model and previous study efforts.

Stakolich, E. G.↗

NASA NTRS2013

Ion Propulsion: An Enabling Technology for the Dawn Mission

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape (heliocentic) trajectory with an escape velocity of 11 km/s. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide an additional delta-V of approximately 11 km/s for the heliocentric transfers to each body and for all orbit transfers including orbit capture/escape and transition to the various science orbits. Deterministic thrusting to Vesta began in December 2007 and concluded with orbit capture at Vesta in July 2011. The transfer to Vesta included a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s and was the only post-launch mission delta-V not provided by IPS. The IPS was used during the 14 months at Vesta for all science orbit transfers and then for Vesta escape. Deterministic thrusting for Ceres began in late August 2012 with a planned arrival date at Ceres in early 2015, whereupon IPS will be used for all science orbit transfers. As of January 2013 the IPS has been operated for approximately 28,000 hours, consumed approximately 280 kg of xenon, and provided a delta-V of approximately 7.5 km/s, the most post-launch delta-V of any spacecraft yet flown. IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. Use of the IPS together with a medium-priced launch vehicle enabled this high delta-V mission to be performed within the Discovery Program cost cap. This paper provides an overview of the Dawn IPS and its mission operations through departure for Ceres.

Dawn mission overview↗

NASA NTRS1988

Development of a liquid-fed water resistojet

A concept for a forced-flow once-through water vaporizer for application to resistojet thrusters was evaluated as an element of a laboratory model thruster and tested to investigate its operating characteristics. The vaporizer design concept employs flow swirling to attach the liquid flow to the boiler chamber wall, providing for separation of the two liquid phases. This vaporizer was modified with a nozzle and a centrally-located heater to facilitate vaporization, superheating, and expansion of the propellant, allowing it to function as a resistojet. Performance was measured at thrust levels ranging from 170 to 360 mN and at power levels ranging from 443 to 192 W. Maximum measured specific impulse was 192 sec.

Morren, W. Earl↗

NASA NTRS1988

Development of a liquid-fed water resistojet

A concept for a forced-flow once-through water vaporizer for application to resistojet thrusters was evaluated as an element of a laboratory model thruster and tested to investigate its operating characteristics. The vaporizer design concept employs flow swirling to attach the liquid flow to the boiler chamber wall, providing for separation of the two liquid phases. This vaporizer was modified with a nozzle and a centrally-located heater to facilitatite vaporization, superheating, and expansion of the propellant, allowing it to function as a resistojet. Performance was measured at thrust levels ranging from 170 to 360 mN and at power levels ranging from 443 to 192 W. Maximum measured specific impulse was 192 sec.

Morren, W. Earl↗

NASA NTRS2016

In-Flight Operation of the Dawn Ion Propulsion System Through the Low Altitude Mapping Orbit at Ceres

The Dawn mission, part of NASA’s Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional ΔV of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to the Approach phase, survey orbit, the high altitude mapping orbit (HAMO), and the low altitude mapping orbit (LAMO) with arrival to LAMO on December 13, 2015, almost eight years after the start of deterministic thrusting to Vesta. The LAMO orbit, at a mean altitude above Ceres of approximately 385 km, is the spacecraft’s final destination and there are no plans to move the spacecraft from LAMO once science operations there are completed. Since arrival at LAMO Dawn’s IPS has been used for occasional orbit maintenance maneuvers while the spacecraft performs scientific investigations. Dawn has successfully completed its science goals and Dawn’s primary mission is scheduled to end June 30, 2016. To date the IPS has been operated for approximately 48,458 hours, consumed approximately 401 kg of xenon, and provided a delta-V of over 11.0 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawn’s IPS continues to be fully operational as of June 2016. This paper provides an overview of Dawn’s mission objectives and the results of Dawn IPS mission operations from Survey orbit through the completion of Dawn’s primary mission.

Garner, Charles E.↗

NASA NTRS2016

In-flight Operation of the Dawn Ion Propulsion System Through the Low Altitude Mapping Orbit at Ceres

The Dawn mission, part of NASA’s Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional ΔV of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to the Approach phase, survey orbit, the high altitude mapping orbit (HAMO), and the low altitude mapping orbit )LAMO) with arrival to LAMO on December 13, 2015, almost eight years after the start of deterministic thrusting to Vesta. The LAMO orbit, at a mean altitude above Ceres of approximately 385 km, is the spacecraft’s final destination and there are no plans to move the spacecraft from LAMO once science operations there are completed. Since arrival at LAMO Dawn’s IPS has been used for occasional orbit maintenance maneuvers while the spacecraft performs scientific investigations. Dawn has successfully completed its science goals and Dawn’s primary mission is scheduled to end in the summer of 2016. To date the IPS has been operated for over 48,454 hours, consumed approximately 401 kg of xenon, and provided a delta-V of over 11.0 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawn’s IPS continues to be fully operational as of June 2016. This paper provides an overview of Dawn’s mission objectives and the results of Dawn IPS mission operations from Survey orbit through June 2016.

Garner, Charles E.↗

NASA NTRS2018

In-Flight Operation of the Dawn Ion Propulsion System Through Completion of the Final Orbit Around Dwarf Planet Ceres

The Dawn mission, part of NASA’s Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional DV of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to all science orbits at Ceres including the final orbit for Dawn’s primary mission, called the low altitude mapping orbit (LAMO), a circular orbit at a mean altitude above Ceres of approximately 385 km. Dawn met or exceeded all pre-launch science requirements and Dawn's prime mission concluded on June 30, 2016. Dawn subsequently received NASA approval for two extended missions at Ceres, called XM1 and XM2. During XM1 IPS was used for orbit maintenance and to transit the spacecraft to several new, higher-altitude science orbits, ending in an elliptical orbit at a maximum altitude of approximately 38,000 km. In XM2, IPS operations included transiting the Dawn spacecraft to its final, highly elliptical orbit at Ceres ranging from 35 km perigee to 4,800 km apogee. Science data acquisition will continue in XM2 until the hydrazine is exhausted, which is expected to occur between August and October 2018. Dawn has successfully completed all science goals for both the primary and XM1 extended missions. To date the IPS has been operated for approximately 51,250 hours, consumed approximately 416 kg of xenon, and provided a delta-V of almost 11.5 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawn’s IPS continues to be fully operational as of June 2018. This paper provides an overview of Dawn’s mission objectives and the results of Dawn IPS mission operations for XM1, and XM2 through June 2018.

Rayman, Marc D.↗