This dissertation is cognizant of the trend toward distributed computing, and the data management... more This dissertation is cognizant of the trend toward distributed computing, and the data management obstacles that must be overcome to achieve efficient, coherent solutions to large-scale finite-element problems. It forms the foundation for a highly structured, object-oriented, data management methodology, which offers portability, scalability of function, and access at multiple levels of abstraction. The methodology is based upon hierarchical multi-level substructuring (MLS) trees and methods. MLS data sets, or structures, are obtained by the decomposition or partitioning of a component. The following topics are presented: (1) Analysis and design of a polynomial time algorithm to approximate an optimal decomposition yielding a hierarchical MLS tree. (2) Design and analysis of a polynomial time algorithm to reorder the equations of a given coefficient matrix in an efficient manner subject to solution constraints. (3) A "tree compression" technique to reduce redundant computa...
The expressions substituted for the costate differential equations are the same for the Earth-cen... more The expressions substituted for the costate differential equations are the same for the Earth-centered and Moon-centered coordinate frames with the states referenced to the respective frames. Since the trajectory terminates in circular orbits, Vr is zero at both ends. Also, as previously indicated, \g is zero in the two terminal circular orbits. Therefore, equations (C.5) and (C.9) determine the position costates and at both ends of the trajectory given the values u, it, v, and v in circular Earth orbit and circular lunar orbit.
Advances in Aerospace Science and Technology, 2017
A new guidance scheme for the approach and landing (A & L) phase of an unpowered reusable launch ... more A new guidance scheme for the approach and landing (A & L) phase of an unpowered reusable launch vehicle (RLV) has been developed. The main advantage of the new guidance is the use of glide-efficiency factor as the independent variable to compute the geometrical flare parameters by a set of analytical functions. The trajectory-planning algorithm generates its reference geometry based on the steep and shallow subphases, respectively. During the steep segment, the quasi-equilibrium glide (QEG) solution, which assumes a constant dynamic pressure and flight-path angle during the flight, is used to create the flight reference while the shallow segment is defined by polynomial functions for altitude and dynamic pressure profiles. Standard linearization methods are used to design a closed-loop command in order to track the QEG profile. Furthermore, proportion-derivative (PD) control is used to modulate the lift coefficient during the flare flight. Once the reference trajectory is created, a closed-loop simulation is obtained to track the reference. Off-nominal conditions, in terms of change in initial glide-efficiency factor, dynamic pressure, flight-path angle, and altitude are tested using a Monte-Carlo simulation. The simulated results demonstrate the effectiveness of the proposed algorithm to land the vehicle successfully under large dispersions of glide-efficiency factor.
Craig Kluever 's Dynamic Systems: Modeling, Simulation, and Control highlights essential topics s... more Craig Kluever 's Dynamic Systems: Modeling, Simulation, and Control highlights essential topics such as analysis, design, and control of physical engineering systems, often composed of interacting mechanical, electrical and fluid subsystem components. The major topics covered in this text include mathematical modeling, system-response analysis, and an introduction to feedback control systems. Dynamic Systems integrates an early introduction to numerical simulation using MATLAB®'s Simulink for integrated systems. Simulink® and MATLAB® tutorials for both software programs will also be provided. The author's text also has a strong emphasis on real-world case studies.
Abstract : This report documents the process, analysis of alternatives and decisions made by an u... more Abstract : This report documents the process, analysis of alternatives and decisions made by an undergraduate team from the University of Missouri-Columbia in designing a radio controlled (RC) aircraft to perform a specific sensor related mission in the 43rd AIAA Joint Propulsion Conference Student Design Challenge. The challenge was to design an integrated propulsion and power system capable of sustaining flight for video surveillance of ground targets while generating additional power to drive a power consuming device Aircraft design constraints included a single propeller vehicle with a maximum take-off weight of 15 lbs; a commercially available airframe with wing span 80-82 in. and fuselage length 62-67 in; a power consuming device with a 28 volt input and integrated into the airframe or mounted externally on the airframe. The amount of power consumption and degree of continuous video surveillance were measures of merit for the design challenge.
AIAA Atmospheric Flight Mechanics Conference and Exhibit, 2004
A guidance scheme that employs a trajectory-planning algorithm has been developed for the termina... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the terminal area energy management phase of an unpowered reusable launch vehicle. The trajectory-planning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. Iterating on three geometric parameters and propagating a test matrix of trajectories produces a set of feasible reference profiles that bring the vehicle from its current state to a desired approach and landing target state. Path constraints on states and controls are easily implemented in the propagation scheme, and simple numerical sorting is used to identify the optimal feasible path. Open- and closed-loop guidance commands are readily available once the best reference trajectory is determined. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in initial vehicle energy, initial heading, and vehicle drag. The effectiveness of the trajectory-planning algorithm is demonstrated by several numerical simulations, which show that the guided vehicle is able to successfully reach the desired approach and landing glideslope target with small tracking errors.
AIAA Guidance, Navigation, and Control Conference and Exhibit, 2005
A guidance scheme that employs a trajectory-planning algorithm has been developed for the termina... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the terminal area energy management phase of an unpowered reusable launch vehicle. The trajectory-planning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. Iterating on three geometric parameters and propagating a test matrix of trajectories produces a set of feasible reference profiles that bring the vehicle from its current state to a desired approach and landing target state. Path constraints on states and controls are easily implemented in the propagation scheme, and simple numerical sorting is used to identify the optimal feasible path. Open- and closed-loop guidance commands are readily available once the best reference trajectory is de termined. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in initial vehicle energy, initial heading, and vehicle drag. The effectiveness of the trajectory-planning algorithm is de monstrated by several numerical simulations. Next, optimal TAEM trajectories are obtained by using an inverse dynamics approach and nonlinear programming methods. These optimal paths exhibit reduced control effort (such as smaller bank angle profiles) when compared to the guided results, and suggest that significant improvements can still be achieved over a guidance method based on a geometric reference path.
AIAA Guidance, Navigation, and Control Conference, 2010
Much effort has been put into developing technologies for next generation re-usable launch vehicl... more Much effort has been put into developing technologies for next generation re-usable launch vehicles. Fully re-usable launch vehicles include a booster stage that is designed to land, usually near the launch site, after it has released the upper-stage, which continues to orbit. The fuel reserve needed to turn the booster stage around will usually be minimal, in order to provide as much energy to the upper stage as possible. For this reason, once the booster stage has completed a rocket-back maneuver, it will typically be at a high altitude (exo-atmospheric) but with a low horizontal velocity and a correspondingly steep flight path angle on re-entry. Traditional re-entry guidance is designed for vehicles with a high velocity, shallow flight path angle, and with strong constraints on heating, and thus these traditional approaches are not appropriate for a low energy re-entry. We have developed a set of guidance algorithms that will successfully guide a vehicle to landing starting from a Low Energy Re-entry (LOER) condition. The guidance algorithms are based on acquiring an Equilibrium Steep Glideslope that ensures the vehicle can achieve near optimal range performance. The guidance problem can be successfully solved for a wide range of initial conditions, including more traditional initial conditions that would occur at the end of a High Energy Re-entry (HIER) from orbit. Thus, the guidance approach we have developed can be used as a more robust version of Terminal Area Energy Management (TAEM) guidance, as well as for LOER and has been tested for a wide range of vehicles, including the Space Shuttle and vehicles with a wide variety of L/D capability.
41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2005
Solar electric propulsion (SEP) is being used for a variety of planetary missions sponsored by ES... more Solar electric propulsion (SEP) is being used for a variety of planetary missions sponsored by ESA, JAXA, and NASA and nuclear electric propulsion (NEP) is being considered for future, flagship-class interplanetary missions. Radioisotope electric propulsion (REP) has recently been shown to effectively complement SEP and NEP for missions to high-AU targets with modest payload requirements. This paper investigates the application of an advanced REP for a sample return from the comet Tempel 1. A set,of mission and system parameters are varied with the goal of quantifying their impact on total mission payload. Mission parameters considered include trip-time and Earth return entry interface speed of the sample return system. System parameters considered include launch vehicle, power level of spacecraft at beginning of mission, and thruster specific impulse. For the baseline case of Atlas 401 and REP power level of 750 W, the mission time was 12 years, the payload was 144 kg, and the missions optimized to a single specific impulse generally within Hall ion thruster range. Other cases were investigated in support of graduate studies, and include the larger Atlas 551 launch vehicle and extended power level to 1 kW. The Atlas 55 1 cases tended to optimize dual specific impulses generally in the Hall ion thruster range for both legs of the mission. A power level of at least I-kW and trip-time of approximately 11 years was required to obtain a total science payload close to 320 kg for the Atlas 401 launch vehicle. An Atlas 551 launch vehicle yielded a science payload of approximately 540 kg for the case of I-kW of power and an 11-year trip time, and nearly 250 kg of science payload for the case of 1-kW of power and a 6-year trip time. Results are also reported indicating the performance ramifications of meeting a reduced Earth entry interface velocity constraint.
Low-thrust geostationary transfer orbits (LT2GEO) are found to offer significant low cost options... more Low-thrust geostationary transfer orbits (LT2GEO) are found to offer significant low cost options thus making them very attractive for GEO missions. However, LT2GEOs increase the transfer orbit time from days to months, thereby causing a significant increase in the time that satellites traverse the most intense trapped particle radiation belts. This paper describes the LT2GEO radiation environment and provides some practical implications using new solar cell technologies and materials. New solar cell and coverglass testing protocols are established using Monte Carlo transport modeling and experimental results are given for Qioptiq CMG borosilicate coverglass.
Maximum-payload transfers to geostationary orbit are computed for solar-electric-propulsion space... more Maximum-payload transfers to geostationary orbit are computed for solar-electric-propulsion spacecraft that employ specific impulse (or equivalently thrust) modulation. The optimal specific impulse profile is obtained by using a direct optimization method that is based on a calculus of variations approach. Engine models with both constant and variable efficiency are considered, and constant input power is assumed. Numerical simulations show that varying specific impulse for a Hall thruster increases the delivered payload mass when compared to transfers with a fixed specific impulse. When thruster efficiency remains constant, modulating specific impulse increases transportation rate (the payload mass advantage over a chemical-propulsion transfer divided by transfer time) by about 5-6%. However, when a more realistic variable-efficiency thruster model is applied, the transportation rate gain from thrust modulation is diminished by a factor of three. This analysis demonstrates that modulating specific impulse would offer little payload delivery gain for geostationary-orbit-raising missions using realistic Hall-thruster models.
Anewdeep-spacemissiontotheheliosphericboundaryisanalyzedforspacecraftutilizinglow-thrustionengine... more Anewdeep-spacemissiontotheheliosphericboundaryisanalyzedforspacecraftutilizinglow-thrustionengines. The mission design is performed by optimizing a combination of trajectory variables and propulsion system parameters such that the total trip time to the heliospheric boundary is minimized. Spacecraft utilizing both solar electricpropulsionandnuclearelectricpropulsionareconsidered.Optimalmissiondesignsarepresentedforawide range of launch vehicle options. In addition, a sensitivity analysis of the assumed electric propulsion technology level and gravity-assist e yby conditions is performed. Although electric propulsion is typically associated with payload fraction enhancement, this analysis demonstrates that the use of low-thrust spacecraft results in relatively short trip times for high-energy deep-space missions. Furthermore, it is shown that the performance of both electric-propulsion spacecraft trajectories compare favorably and in many cases show improvement overmissions utilizing all-chemical propulsion systems.
ABSTRACT A new combined vehicle-and-trajectory optimization problem is solved for a low-thrust nu... more ABSTRACT A new combined vehicle-and-trajectory optimization problem is solved for a low-thrust nuclear-electric-propulsion spacecraft whose motion is governed by restricted three-body-problem dynamics for the earth-moon system. The problem involves computing the optimal spacecraft sizing parameters and trajectory design variables that result in the maximum payload for a fixed-trip-time, planar transfer from circular low earth orbit to circular low-lunar orbit. In particular, the optimal specific impulse and input power are computed. The detailed vehicle-and-trajectory optimization approach is effective in solving a complex interactive problem that is important to both spacecraft and mission designers. Several numerical solutions are obtained for a wide range of trip times.
increased service bandwidth. The Intelsat series of ABSTRACT satellites present a good example of... more increased service bandwidth. The Intelsat series of ABSTRACT satellites present a good example of these trends. 1 Solar Electric Propulsion (SEP) technology is currently Intelsat 1 and 2, launched during the late sixties, had being used for geostationary satellite station keeping to lifetimes under four years. Intelsats 4 and 5 had seven increase payload mass. Analyses show that advanced year design lifetimes. Intelsat 7 had full capacity design electric propulsion technologies can be used to obtain lifetime of ten years with propellant for 15 years. The additional increases in payload mass by using these same planned Intelsat 8/8A series lifetime is 14-18 years using technologies to perform part of the orbit transfer. In this N2H4 arcjets for station keeping. These results indicate a work three electric propulsion technologies are examined continuing trend toward longer lifetimes, thus a 15 year at two power levels for an Atlas IIAS class spacecraft, lifetime is assumed in these analyses. Satellite masses, The on-board chemical propulsion apogee engine fuel is and the launch vehicles to deliver them, have also gown. reduced in this analysis to allow the use of electric Early Intelsats were well under 1000 kg dry mass. The propulsion. A numerical optimizer is used to determine planned Intelsat 8/8A series will have a 1530kg dry mass. the chemical burns which will minimize the electric End-of-life (EOL) power levels have increased from propulsion transfer time. Results show that for a 1550 kg hundredsof watts for Intelsats 1 to 4, to over 5 kW for Atlas gAS class payload, increases in net mass Intelsat 7A. Intelsat 8/8A will use the Martin Marietta (geostationary satellite mass less wet propulsion system Astro Space Series 7000 which has a beginning of life mass) of 150 to 800 kg are possible using electric (BOL) power level over 7 kW. Finally, communication propulsion for station keeping, advancedchemical engines bandwidthson Intelsat spacecraft have increased from 50 for part of the transfer and electric propulsion for the MHz on Intelsat 1 to 2856 MHz on the planned Intelsat remainderof the transfer. Trip times are between one and 8/8A series. These continuing trends toward larger, more four months, capable, longer life and higher power spacecraft were used to select the spacecraft characteristics in this study. Higher INTRODUCTION power spacecraft permit expansion of the use of electric Solar Electric Propulsion (SEP) is already being used for propulsion systems beyond the already demonstrated station keeping of geostationary satellites, most notably station keeping function to encompass a portion of the hydrazine arcjets on AT&T's Telstar 4 and SPT-'100Hall orbit transfer mission. Successful implementation of thrusters on the Russian GALS spacecraft.1 The next step advanced propulsion systems will enable continued gowth in the development of electric propulsion systems is to use of geostationary satellite capability without requiring these types of thrusters to contribute to placing the growth in spacecraft mass or launch vehicle and will • spacecraft into geostationary orbit. For a given launch permit continued expansion of communicationscapability. vehicle, the fuel mass savings could then be directly used to increase the payload, for instance, the number of Studies by various authors have shown the net mass communication transponders. Even a small increase in benefits of using electric propulsion for transfer from mass might have large revenue impacts, various high Earth orbits2,3,Appendix in order to avoid the long trip times and Van Alien belt radiation damage of The current trend for geostationary spacecraft is towards low Earth orbit (LEO) to geostationary Earth orbit (GEO) longer lifetimes, increased masses, higher powers, and transfers using electric propulsion. 4,5 However, none of This paper is declared a work of the U.S. Government and 1 is not subject to copyright protection in the United States.
A guidance scheme that employs a trajectory-planning algorithm has been developed for the approac... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the approach and landing phase of an unpowered reusable launch vehicle. The trajectoryplanning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. A feasible reference profile that brings the vehicle from its current state to a desired landing condition is obtained by iterating on a single geometric parameter, and the flight-path angle at the start of the flare is selected as the iteration variable. Open- and closed-loop guidance commands are readily available once the reference trajectory is obtained. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in winds, vehicle energy, and vehicle drag. The effectiveness of the tr ajectoryplanning algorithm is demonstrated by several numerical simulations, which show that the guided vehicle is able to land successfully with adequate energy margin.
This dissertation is cognizant of the trend toward distributed computing, and the data management... more This dissertation is cognizant of the trend toward distributed computing, and the data management obstacles that must be overcome to achieve efficient, coherent solutions to large-scale finite-element problems. It forms the foundation for a highly structured, object-oriented, data management methodology, which offers portability, scalability of function, and access at multiple levels of abstraction. The methodology is based upon hierarchical multi-level substructuring (MLS) trees and methods. MLS data sets, or structures, are obtained by the decomposition or partitioning of a component. The following topics are presented: (1) Analysis and design of a polynomial time algorithm to approximate an optimal decomposition yielding a hierarchical MLS tree. (2) Design and analysis of a polynomial time algorithm to reorder the equations of a given coefficient matrix in an efficient manner subject to solution constraints. (3) A "tree compression" technique to reduce redundant computa...
The expressions substituted for the costate differential equations are the same for the Earth-cen... more The expressions substituted for the costate differential equations are the same for the Earth-centered and Moon-centered coordinate frames with the states referenced to the respective frames. Since the trajectory terminates in circular orbits, Vr is zero at both ends. Also, as previously indicated, \g is zero in the two terminal circular orbits. Therefore, equations (C.5) and (C.9) determine the position costates and at both ends of the trajectory given the values u, it, v, and v in circular Earth orbit and circular lunar orbit.
Advances in Aerospace Science and Technology, 2017
A new guidance scheme for the approach and landing (A & L) phase of an unpowered reusable launch ... more A new guidance scheme for the approach and landing (A & L) phase of an unpowered reusable launch vehicle (RLV) has been developed. The main advantage of the new guidance is the use of glide-efficiency factor as the independent variable to compute the geometrical flare parameters by a set of analytical functions. The trajectory-planning algorithm generates its reference geometry based on the steep and shallow subphases, respectively. During the steep segment, the quasi-equilibrium glide (QEG) solution, which assumes a constant dynamic pressure and flight-path angle during the flight, is used to create the flight reference while the shallow segment is defined by polynomial functions for altitude and dynamic pressure profiles. Standard linearization methods are used to design a closed-loop command in order to track the QEG profile. Furthermore, proportion-derivative (PD) control is used to modulate the lift coefficient during the flare flight. Once the reference trajectory is created, a closed-loop simulation is obtained to track the reference. Off-nominal conditions, in terms of change in initial glide-efficiency factor, dynamic pressure, flight-path angle, and altitude are tested using a Monte-Carlo simulation. The simulated results demonstrate the effectiveness of the proposed algorithm to land the vehicle successfully under large dispersions of glide-efficiency factor.
Craig Kluever 's Dynamic Systems: Modeling, Simulation, and Control highlights essential topics s... more Craig Kluever 's Dynamic Systems: Modeling, Simulation, and Control highlights essential topics such as analysis, design, and control of physical engineering systems, often composed of interacting mechanical, electrical and fluid subsystem components. The major topics covered in this text include mathematical modeling, system-response analysis, and an introduction to feedback control systems. Dynamic Systems integrates an early introduction to numerical simulation using MATLAB®'s Simulink for integrated systems. Simulink® and MATLAB® tutorials for both software programs will also be provided. The author's text also has a strong emphasis on real-world case studies.
Abstract : This report documents the process, analysis of alternatives and decisions made by an u... more Abstract : This report documents the process, analysis of alternatives and decisions made by an undergraduate team from the University of Missouri-Columbia in designing a radio controlled (RC) aircraft to perform a specific sensor related mission in the 43rd AIAA Joint Propulsion Conference Student Design Challenge. The challenge was to design an integrated propulsion and power system capable of sustaining flight for video surveillance of ground targets while generating additional power to drive a power consuming device Aircraft design constraints included a single propeller vehicle with a maximum take-off weight of 15 lbs; a commercially available airframe with wing span 80-82 in. and fuselage length 62-67 in; a power consuming device with a 28 volt input and integrated into the airframe or mounted externally on the airframe. The amount of power consumption and degree of continuous video surveillance were measures of merit for the design challenge.
AIAA Atmospheric Flight Mechanics Conference and Exhibit, 2004
A guidance scheme that employs a trajectory-planning algorithm has been developed for the termina... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the terminal area energy management phase of an unpowered reusable launch vehicle. The trajectory-planning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. Iterating on three geometric parameters and propagating a test matrix of trajectories produces a set of feasible reference profiles that bring the vehicle from its current state to a desired approach and landing target state. Path constraints on states and controls are easily implemented in the propagation scheme, and simple numerical sorting is used to identify the optimal feasible path. Open- and closed-loop guidance commands are readily available once the best reference trajectory is determined. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in initial vehicle energy, initial heading, and vehicle drag. The effectiveness of the trajectory-planning algorithm is demonstrated by several numerical simulations, which show that the guided vehicle is able to successfully reach the desired approach and landing glideslope target with small tracking errors.
AIAA Guidance, Navigation, and Control Conference and Exhibit, 2005
A guidance scheme that employs a trajectory-planning algorithm has been developed for the termina... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the terminal area energy management phase of an unpowered reusable launch vehicle. The trajectory-planning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. Iterating on three geometric parameters and propagating a test matrix of trajectories produces a set of feasible reference profiles that bring the vehicle from its current state to a desired approach and landing target state. Path constraints on states and controls are easily implemented in the propagation scheme, and simple numerical sorting is used to identify the optimal feasible path. Open- and closed-loop guidance commands are readily available once the best reference trajectory is de termined. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in initial vehicle energy, initial heading, and vehicle drag. The effectiveness of the trajectory-planning algorithm is de monstrated by several numerical simulations. Next, optimal TAEM trajectories are obtained by using an inverse dynamics approach and nonlinear programming methods. These optimal paths exhibit reduced control effort (such as smaller bank angle profiles) when compared to the guided results, and suggest that significant improvements can still be achieved over a guidance method based on a geometric reference path.
AIAA Guidance, Navigation, and Control Conference, 2010
Much effort has been put into developing technologies for next generation re-usable launch vehicl... more Much effort has been put into developing technologies for next generation re-usable launch vehicles. Fully re-usable launch vehicles include a booster stage that is designed to land, usually near the launch site, after it has released the upper-stage, which continues to orbit. The fuel reserve needed to turn the booster stage around will usually be minimal, in order to provide as much energy to the upper stage as possible. For this reason, once the booster stage has completed a rocket-back maneuver, it will typically be at a high altitude (exo-atmospheric) but with a low horizontal velocity and a correspondingly steep flight path angle on re-entry. Traditional re-entry guidance is designed for vehicles with a high velocity, shallow flight path angle, and with strong constraints on heating, and thus these traditional approaches are not appropriate for a low energy re-entry. We have developed a set of guidance algorithms that will successfully guide a vehicle to landing starting from a Low Energy Re-entry (LOER) condition. The guidance algorithms are based on acquiring an Equilibrium Steep Glideslope that ensures the vehicle can achieve near optimal range performance. The guidance problem can be successfully solved for a wide range of initial conditions, including more traditional initial conditions that would occur at the end of a High Energy Re-entry (HIER) from orbit. Thus, the guidance approach we have developed can be used as a more robust version of Terminal Area Energy Management (TAEM) guidance, as well as for LOER and has been tested for a wide range of vehicles, including the Space Shuttle and vehicles with a wide variety of L/D capability.
41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2005
Solar electric propulsion (SEP) is being used for a variety of planetary missions sponsored by ES... more Solar electric propulsion (SEP) is being used for a variety of planetary missions sponsored by ESA, JAXA, and NASA and nuclear electric propulsion (NEP) is being considered for future, flagship-class interplanetary missions. Radioisotope electric propulsion (REP) has recently been shown to effectively complement SEP and NEP for missions to high-AU targets with modest payload requirements. This paper investigates the application of an advanced REP for a sample return from the comet Tempel 1. A set,of mission and system parameters are varied with the goal of quantifying their impact on total mission payload. Mission parameters considered include trip-time and Earth return entry interface speed of the sample return system. System parameters considered include launch vehicle, power level of spacecraft at beginning of mission, and thruster specific impulse. For the baseline case of Atlas 401 and REP power level of 750 W, the mission time was 12 years, the payload was 144 kg, and the missions optimized to a single specific impulse generally within Hall ion thruster range. Other cases were investigated in support of graduate studies, and include the larger Atlas 551 launch vehicle and extended power level to 1 kW. The Atlas 55 1 cases tended to optimize dual specific impulses generally in the Hall ion thruster range for both legs of the mission. A power level of at least I-kW and trip-time of approximately 11 years was required to obtain a total science payload close to 320 kg for the Atlas 401 launch vehicle. An Atlas 551 launch vehicle yielded a science payload of approximately 540 kg for the case of I-kW of power and an 11-year trip time, and nearly 250 kg of science payload for the case of 1-kW of power and a 6-year trip time. Results are also reported indicating the performance ramifications of meeting a reduced Earth entry interface velocity constraint.
Low-thrust geostationary transfer orbits (LT2GEO) are found to offer significant low cost options... more Low-thrust geostationary transfer orbits (LT2GEO) are found to offer significant low cost options thus making them very attractive for GEO missions. However, LT2GEOs increase the transfer orbit time from days to months, thereby causing a significant increase in the time that satellites traverse the most intense trapped particle radiation belts. This paper describes the LT2GEO radiation environment and provides some practical implications using new solar cell technologies and materials. New solar cell and coverglass testing protocols are established using Monte Carlo transport modeling and experimental results are given for Qioptiq CMG borosilicate coverglass.
Maximum-payload transfers to geostationary orbit are computed for solar-electric-propulsion space... more Maximum-payload transfers to geostationary orbit are computed for solar-electric-propulsion spacecraft that employ specific impulse (or equivalently thrust) modulation. The optimal specific impulse profile is obtained by using a direct optimization method that is based on a calculus of variations approach. Engine models with both constant and variable efficiency are considered, and constant input power is assumed. Numerical simulations show that varying specific impulse for a Hall thruster increases the delivered payload mass when compared to transfers with a fixed specific impulse. When thruster efficiency remains constant, modulating specific impulse increases transportation rate (the payload mass advantage over a chemical-propulsion transfer divided by transfer time) by about 5-6%. However, when a more realistic variable-efficiency thruster model is applied, the transportation rate gain from thrust modulation is diminished by a factor of three. This analysis demonstrates that modulating specific impulse would offer little payload delivery gain for geostationary-orbit-raising missions using realistic Hall-thruster models.
Anewdeep-spacemissiontotheheliosphericboundaryisanalyzedforspacecraftutilizinglow-thrustionengine... more Anewdeep-spacemissiontotheheliosphericboundaryisanalyzedforspacecraftutilizinglow-thrustionengines. The mission design is performed by optimizing a combination of trajectory variables and propulsion system parameters such that the total trip time to the heliospheric boundary is minimized. Spacecraft utilizing both solar electricpropulsionandnuclearelectricpropulsionareconsidered.Optimalmissiondesignsarepresentedforawide range of launch vehicle options. In addition, a sensitivity analysis of the assumed electric propulsion technology level and gravity-assist e yby conditions is performed. Although electric propulsion is typically associated with payload fraction enhancement, this analysis demonstrates that the use of low-thrust spacecraft results in relatively short trip times for high-energy deep-space missions. Furthermore, it is shown that the performance of both electric-propulsion spacecraft trajectories compare favorably and in many cases show improvement overmissions utilizing all-chemical propulsion systems.
ABSTRACT A new combined vehicle-and-trajectory optimization problem is solved for a low-thrust nu... more ABSTRACT A new combined vehicle-and-trajectory optimization problem is solved for a low-thrust nuclear-electric-propulsion spacecraft whose motion is governed by restricted three-body-problem dynamics for the earth-moon system. The problem involves computing the optimal spacecraft sizing parameters and trajectory design variables that result in the maximum payload for a fixed-trip-time, planar transfer from circular low earth orbit to circular low-lunar orbit. In particular, the optimal specific impulse and input power are computed. The detailed vehicle-and-trajectory optimization approach is effective in solving a complex interactive problem that is important to both spacecraft and mission designers. Several numerical solutions are obtained for a wide range of trip times.
increased service bandwidth. The Intelsat series of ABSTRACT satellites present a good example of... more increased service bandwidth. The Intelsat series of ABSTRACT satellites present a good example of these trends. 1 Solar Electric Propulsion (SEP) technology is currently Intelsat 1 and 2, launched during the late sixties, had being used for geostationary satellite station keeping to lifetimes under four years. Intelsats 4 and 5 had seven increase payload mass. Analyses show that advanced year design lifetimes. Intelsat 7 had full capacity design electric propulsion technologies can be used to obtain lifetime of ten years with propellant for 15 years. The additional increases in payload mass by using these same planned Intelsat 8/8A series lifetime is 14-18 years using technologies to perform part of the orbit transfer. In this N2H4 arcjets for station keeping. These results indicate a work three electric propulsion technologies are examined continuing trend toward longer lifetimes, thus a 15 year at two power levels for an Atlas IIAS class spacecraft, lifetime is assumed in these analyses. Satellite masses, The on-board chemical propulsion apogee engine fuel is and the launch vehicles to deliver them, have also gown. reduced in this analysis to allow the use of electric Early Intelsats were well under 1000 kg dry mass. The propulsion. A numerical optimizer is used to determine planned Intelsat 8/8A series will have a 1530kg dry mass. the chemical burns which will minimize the electric End-of-life (EOL) power levels have increased from propulsion transfer time. Results show that for a 1550 kg hundredsof watts for Intelsats 1 to 4, to over 5 kW for Atlas gAS class payload, increases in net mass Intelsat 7A. Intelsat 8/8A will use the Martin Marietta (geostationary satellite mass less wet propulsion system Astro Space Series 7000 which has a beginning of life mass) of 150 to 800 kg are possible using electric (BOL) power level over 7 kW. Finally, communication propulsion for station keeping, advancedchemical engines bandwidthson Intelsat spacecraft have increased from 50 for part of the transfer and electric propulsion for the MHz on Intelsat 1 to 2856 MHz on the planned Intelsat remainderof the transfer. Trip times are between one and 8/8A series. These continuing trends toward larger, more four months, capable, longer life and higher power spacecraft were used to select the spacecraft characteristics in this study. Higher INTRODUCTION power spacecraft permit expansion of the use of electric Solar Electric Propulsion (SEP) is already being used for propulsion systems beyond the already demonstrated station keeping of geostationary satellites, most notably station keeping function to encompass a portion of the hydrazine arcjets on AT&T's Telstar 4 and SPT-'100Hall orbit transfer mission. Successful implementation of thrusters on the Russian GALS spacecraft.1 The next step advanced propulsion systems will enable continued gowth in the development of electric propulsion systems is to use of geostationary satellite capability without requiring these types of thrusters to contribute to placing the growth in spacecraft mass or launch vehicle and will • spacecraft into geostationary orbit. For a given launch permit continued expansion of communicationscapability. vehicle, the fuel mass savings could then be directly used to increase the payload, for instance, the number of Studies by various authors have shown the net mass communication transponders. Even a small increase in benefits of using electric propulsion for transfer from mass might have large revenue impacts, various high Earth orbits2,3,Appendix in order to avoid the long trip times and Van Alien belt radiation damage of The current trend for geostationary spacecraft is towards low Earth orbit (LEO) to geostationary Earth orbit (GEO) longer lifetimes, increased masses, higher powers, and transfers using electric propulsion. 4,5 However, none of This paper is declared a work of the U.S. Government and 1 is not subject to copyright protection in the United States.
A guidance scheme that employs a trajectory-planning algorithm has been developed for the approac... more A guidance scheme that employs a trajectory-planning algorithm has been developed for the approach and landing phase of an unpowered reusable launch vehicle. The trajectoryplanning scheme computes a reference flight profile by piecing together several flight segments that are defined by a small set of geometric parameters. A feasible reference profile that brings the vehicle from its current state to a desired landing condition is obtained by iterating on a single geometric parameter, and the flight-path angle at the start of the flare is selected as the iteration variable. Open- and closed-loop guidance commands are readily available once the reference trajectory is obtained. The trajectory-planning algorithm is able to quickly generate new reference profiles for test cases with large variations in winds, vehicle energy, and vehicle drag. The effectiveness of the tr ajectoryplanning algorithm is demonstrated by several numerical simulations, which show that the guided vehicle is able to land successfully with adequate energy margin.
Uploads
Papers by Craig Kluever