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Andrew D. Ketsdever - One of the best experts on this subject based on the ideXlab platform.
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Development of a Specific Impulse Balance for Capillary Discharge Pulsed Plasma Thrusters
Journal of Propulsion and Power, 2009Co-Authors: Taylor C. Lilly, Andrew D. Ketsdever, Anthony P. Pancotti, Marcus YoungAbstract:tot mpropg0 (1) where Itot is the total impulse, mprop is the total Propellant Mass loss, and g0 is the gravitational constant. The thruster will be configured on the thrust stand such that the impulse generated by the discharge and the steady-state force generated by the Propellant Mass loss act in the same direction. The combined signal from these effects can then be decoupled to assess the ratio of the impulse to the weight of Propellant expended, yielding the specific impulse.
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Thrust Stand MicroMass Balance for the Direct Measurement of Specific Impulse
Journal of Propulsion and Power, 2008Co-Authors: Andrew D. Ketsdever, Brian D'souza, Riki H. LeeAbstract:A technique has been developed to directly measure the specific impulse from pulsed thruster systems. The technique is especially useful for propulsion devices that use solid Propellants, for which a direct measurement of the Propellant Mass flow is extremely difficult. A torsion balance is used with a horizontal axis of rotation. A thruster is placed on the balance such that the impulse of the thruster firing and the change in Mass due to the expelled Propellant act in the same direction. A combined impulse and steady-state force measurement (due to Propellant Mass loss) can then be decoupled to assess the ratio of the impulse to the weight of Propellant expended, or the specific impulse. A model has been developed to show the utility of the technique for pulsed systems with a firing time less than the natural period of the balance. An experimental proof of principle study was also undertaken using the laser ablation of engineering-grade Buna, Viton, and Teflon Propellants. Specific-impulse measurements on the order of 200 s have been demonstrated with this laser ablation thruster.
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Development of a Specific Impulse Balance for a Pulsed Capillary Discharge (Preprint)
2008Co-Authors: Andrew D. Ketsdever, Taylor C. Lilly, Anthony P. Pancotti, Marcus Young, J A DuncanAbstract:Abstract : A capillary discharge based pulsed plasma thruster is currently under development at the Air Force Research Laboratory. A torsion thrust stand has been developed to simultaneously measure the shot by shot total impulse and Propellant Mass loss of the thruster, yielding a per shot specific impulse. Competing design considerations affecting the performance of the thrust stand has lead to an optimized diagnostic tool capable of measuring total impulses up to 150 mN s with a resolution of 0.4 mN s and Mass losses up to 18g with a resolution of 1 mg.
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Thrust Stand Mass Balance Measurements of Hybrid Motor Mass Flow
43rd AIAA ASME SAE ASEE Joint Propulsion Conference & Exhibit, 2007Co-Authors: J. D. Olliges, Taylor C. Lilly, Miles Killingsworth, Andrew D. KetsdeverAbstract:A novel diagnostic technique has been developed, utilizing the Thrust Stand Mass Balance, to directly measure a time accurate Mass flow from a solid-fuel thruster. The Mass flow measurement technique has been verified using an idealized numerical simulation. Two calibration experiments have been performed to assess the dynamic response of the Mass balance. First, a set of calibration weights were placed on the Mass balance and removed in order to properly characterize the Mass balance motion. Second, a known Mass flow rate of water was deposited onto the test stand. As a proof of concept experiment, a 3.81cm diameter PMMA/GOx hybrid thruster core was burned and the Propellant Mass flow was measured. Variations in the GOx flow rate resulted in corresponding variations in the total Propellant Mass flow as expected, showing the utility of the Thrust Stand Mass Balance as a Mass flow measurement device.
Taylor C. Lilly - One of the best experts on this subject based on the ideXlab platform.
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Development of a Specific Impulse Balance for Capillary Discharge Pulsed Plasma Thrusters
Journal of Propulsion and Power, 2009Co-Authors: Taylor C. Lilly, Andrew D. Ketsdever, Anthony P. Pancotti, Marcus YoungAbstract:tot mpropg0 (1) where Itot is the total impulse, mprop is the total Propellant Mass loss, and g0 is the gravitational constant. The thruster will be configured on the thrust stand such that the impulse generated by the discharge and the steady-state force generated by the Propellant Mass loss act in the same direction. The combined signal from these effects can then be decoupled to assess the ratio of the impulse to the weight of Propellant expended, yielding the specific impulse.
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Development of a Specific Impulse Balance for a Pulsed Capillary Discharge (Preprint)
2008Co-Authors: Andrew D. Ketsdever, Taylor C. Lilly, Anthony P. Pancotti, Marcus Young, J A DuncanAbstract:Abstract : A capillary discharge based pulsed plasma thruster is currently under development at the Air Force Research Laboratory. A torsion thrust stand has been developed to simultaneously measure the shot by shot total impulse and Propellant Mass loss of the thruster, yielding a per shot specific impulse. Competing design considerations affecting the performance of the thrust stand has lead to an optimized diagnostic tool capable of measuring total impulses up to 150 mN s with a resolution of 0.4 mN s and Mass losses up to 18g with a resolution of 1 mg.
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Thrust Stand Mass Balance Measurements of Hybrid Motor Mass Flow
43rd AIAA ASME SAE ASEE Joint Propulsion Conference & Exhibit, 2007Co-Authors: J. D. Olliges, Taylor C. Lilly, Miles Killingsworth, Andrew D. KetsdeverAbstract:A novel diagnostic technique has been developed, utilizing the Thrust Stand Mass Balance, to directly measure a time accurate Mass flow from a solid-fuel thruster. The Mass flow measurement technique has been verified using an idealized numerical simulation. Two calibration experiments have been performed to assess the dynamic response of the Mass balance. First, a set of calibration weights were placed on the Mass balance and removed in order to properly characterize the Mass balance motion. Second, a known Mass flow rate of water was deposited onto the test stand. As a proof of concept experiment, a 3.81cm diameter PMMA/GOx hybrid thruster core was burned and the Propellant Mass flow was measured. Variations in the GOx flow rate resulted in corresponding variations in the total Propellant Mass flow as expected, showing the utility of the Thrust Stand Mass Balance as a Mass flow measurement device.
J. D. Olliges - One of the best experts on this subject based on the ideXlab platform.
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Thrust Stand Mass Balance Measurements of Hybrid Motor Mass Flow
43rd AIAA ASME SAE ASEE Joint Propulsion Conference & Exhibit, 2007Co-Authors: J. D. Olliges, Taylor C. Lilly, Miles Killingsworth, Andrew D. KetsdeverAbstract:A novel diagnostic technique has been developed, utilizing the Thrust Stand Mass Balance, to directly measure a time accurate Mass flow from a solid-fuel thruster. The Mass flow measurement technique has been verified using an idealized numerical simulation. Two calibration experiments have been performed to assess the dynamic response of the Mass balance. First, a set of calibration weights were placed on the Mass balance and removed in order to properly characterize the Mass balance motion. Second, a known Mass flow rate of water was deposited onto the test stand. As a proof of concept experiment, a 3.81cm diameter PMMA/GOx hybrid thruster core was burned and the Propellant Mass flow was measured. Variations in the GOx flow rate resulted in corresponding variations in the total Propellant Mass flow as expected, showing the utility of the Thrust Stand Mass Balance as a Mass flow measurement device.
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Thrust Stand Mass Balance Measurements of Hybrid Motor Mass Flow (Postprint)
2007Co-Authors: J. D. Olliges, Miles Killingsworth, T. C. Lilly, A. D. KetsdeverAbstract:Abstract : A novel diagnostic technique has been developed, utilizing the Thrust Stand Mass Balance, to directly measure a time accurate Mass flow from a solid-fuel thruster for systems where the Mass flow rate is of the same order as the experimental error. The Mass flow measurement technique has been verified using an idealized numerical simulation. Two calibration experiments have been performed to assess the dynamic response of the Mass balance. First, a set of calibration weights were placed on the Mass balance and removed in order to properly characterize the Mass balance motion. Second, a known Mass flow rate of water was deposited onto the test stand. As a proof of concept experiment, a 3.81cm diameter PMMA/GOx hybrid thruster core was burned and the Propellant Mass flow was measured. Variations in the GOx flow rate resulted in corresponding variations in the total Propellant Mass flow as expected, showing the utility of the Thrust Stand Mass Balance as a Mass flow measurement device.
Paul Von Allmen - One of the best experts on this subject based on the ideXlab platform.
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Low-thrust orbit transfer optimization with refined Q-law and multi-objective genetic algorithm
2005Co-Authors: Seungwon Lee, Anastassios E. Petropoulos, Paul Von AllmenAbstract:An optimization method for low-thrust orbit transfers around a central body is developed using the Q-law and a multi-objective genetic algorithm. in the hybrid method, the Q-law generates candidate orbit transfers, and the multi-objective genetic algorithm optimizes the Q-law control parameters in order to simultaneously minimize both the consumed Propellant Mass and flight time of the orbit tranfer. This paper addresses the problem of finding optimal orbit transfers for low-thrust spacecraft.
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Multi-Objective Evolutionary Algorithms for Low-Thrust Orbit Transfer Optimization
2005Co-Authors: Seungwon Lee, Paul Von Allmen, Wolfgang Fink, Anastassios E. Petropoulos, Richard J. TerrileAbstract:We address the problem of optimizing a spacecraft trajectory by using three different multi-objective evolutionary algorithms: i) Non-dominated sorting genetic algorithm, ii) Pareto-based ranking genetic algorithm, and iii) Strength Pareto genetic algorithm. The trajectory of interest is an orbit transfer around a central body when the spacecraft uses a lowthrust propulsion system. We use a Lyapunov feedback control law called the Q-law to create an eligible trajectory, while the Q-law control parameters are selected with the multiobjective algorithms. The optimization goal is to minimize flight time and consumed Propellant Mass simultaneously. The Pareto fronts (trade-off surface between flight time and Propellant Mass) produced by these algorithms are evaluated by means of two quantitative metrics: 1) size of the dominated space and 2) coverage of two Pareto fronts. With the two metrics, a hierarchy of algorithms emerged. The nondominated sorting genetic algorithm and the strength Pareto genetic algorithm are equally effective, and they outperform the Pareto-based ranking genetic algorithm.
Angelo Miele - One of the best experts on this subject based on the ideXlab platform.
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Computation of optimal Mars trajectories via combined chemical/electrical propulsion, Part 3: Compromise solutions ☆
Acta Astronautica, 2005Co-Authors: Angelo Miele, T. Wang, P. N. WilliamsAbstract:Abstract The success of the solar-electric ion engine powering the DS1 spacecraft has paved the way toward the use of low-thrust electrical engines in future planetary/interplanetary missions. Vis-a-vis a chemical engine, an electrical engine has a higher specific impulse, implying a possible decrease in Propellant Mass; however, the low-thrust aspect discourages the use of an electrical engine in the near-planet phases of a trip, since this might result in an increase in flight time. Therefore, a fundamental design problem is to find the best combination of chemical propulsion and electrical propulsion for a given mission, for example, a mission from Earth to Mars. With this in mind, this paper is the third of a series dealing with the optimization of Earth–Mars missions via the use of hybrid engines, namely the combination of high-thrust chemical engines for planetary flight and low-thrust electrical engines for interplanetary flight. We look at the deep-space interplanetary portion of the trajectory under rather idealized conditions. The two major performance indexes, the Propellant Mass and the flight time, are in conflict with one another for the following reason: any attempt at reducing the former causes an increase in the latter and vice versa. Therefore, it is natural to consider a compromise performance index involving the scaled values of the Propellant Mass and flight time weighted respectively by the compromise factor C and its complement 1 - C . We use the compromise factor as the parameter of the one-parameter family of compromise trajectories. Analyses carried out with the sequential gradient-restoration algorithm for optimal control problems lead to results which can be highlighted as follows. Thrust profile. Generally speaking, the thrust profile of the compromise trajectory includes three subarcs: the first subarc is characterized by maximum thrust in conjunction with positive (upward) thrust direction; the second subarc is characterized by zero thrust (coasting flight); the third subarc is characterized by maximum thrust in conjunction with negative (downward) thrust direction. Effect of the compromise factor. As the compromise factor increases, the Propellant Mass decreases and the flight time increases; correspondingly, the following changes in the thrust profile take place: (a) the time lengths of the first and third subarcs (powered phases) decrease slightly, meaning that thrust application occurs for shorter duration; also, the average value of the thrust direction in the first and third subarcs decreases, implying higher efficiency of thrust application wrt the spacecraft energy level; as a result, the total Propellant Mass decreases; (b) the time length of the second subarc (coasting) increases considerably, resulting in total time increase. Minimum time trajectory. If C = 0 , the resulting minimum time trajectory has the following characteristics: (a) the time length of the coasting subarc reduces to zero and the three-subarc trajectory degenerates into a two-subarc trajectory; (b) maximum thrust is applied at all times and the thrust direction switches from upward to downward at midcourse. Minimum Propellant Mass trajectory. If C = 1 , the resulting minimum Propellant Mass trajectory has the following characteristics: (a) the thrust magnitude has a bang-zero-bang profile; (b) for the powered subarcs, the thrust direction is tangent to the flight path at all times.
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Optimal Interplanetary Orbital Transfers via Electrical Engines
Journal of Optimization Theory and Applications, 2005Co-Authors: Angelo Miele, T. Wang, P. N. WilliamsAbstract:The Hohmann transfer theory, developed in the 19th century, is the kernel of orbital transfer with minimum Propellant Mass by means of chemical engines. The success of the Deep Space 1 spacecraft has paved the way toward using advanced electrical engines in space. While chemical engines are characterized by high thrust and low specific impulse, electrical engines are characterized by low thrust and hight specific impulse. In this paper, we focus on four issues of optimal interplanetary transfer for a spacecraft powered by an electrical engine controlled via the thrust direction and thrust setting: (a) trajectories of compromise between transfer time and Propellant Mass, (b) trajectories of minimum time, (c) trajectories of minimum Propellant Mass, and (d) relations with the Hohmann transfer trajectory. The resulting fundamental properties are as follows:
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Fundamental Properties of Optimal Orbital Transfers
54th International Astronautical Congress of the International Astronautical Federation the International Academy of Astronautics and the Internationa, 2003Co-Authors: Angelo MieleAbstract:The Hohmann transfer theory, developed in 19th century, is the kernel of orbital transfer with minimum Propellant Mass by means of chemical engines. The success of the Deep Space 1 spacecraft has paved the way toward using advanced electrical engines in space. While chemical engines are characterized by high thrust and low specific impulse, electrical engines are characterized by low thrust and high specific impulse. In this paper, we focus on three issues of optimal orbital transfers for a spacecraft controlled via thrust direction and setting: (a) trajectories of compromise between flight time and Propellant Mass, (b) trajectories of minimum Propellant Mass, and (c) relations with the Hohmann transfer trajectory. The resulting fundamental properties are as follows: (a) Flight Time/Propellant Mass Compromise. For interplanetary orbital transfer (orbital period of order year), an important objective of trajectory optimization is a compromise between flight time and Propellant Mass. The resulting trajectories have a three-subarc thrust profile: the first and third subarcs are characterized by maximum thrust; the second subarc is characterized by zero thrust (coasting flight); for the first subarc, the normal component of the thrust is opposite to that of the third subarc. When the compromise factor shifts from flight time toward Propellant Mass, the average magnitude of the thrust angle for the first and third subarcs decreases, while the flight 1 The development of the multiple-subarc sequential gradient-restoration algorithm employed in this work was supported by NSF Grant CMS-0218878.
