Alan Wilhite | Georgia Institute of Technology (original) (raw)

Papers by Alan Wilhite

Research paper thumbnail of ESAS-Derived Earth Departure Stage Design for Human Mars Exploration

Retrieved …, 2007

Background The Vision for Space Exploration mandated the goal of returning humans to the Moon by ... more Background The Vision for Space Exploration mandated the goal of returning humans to the Moon by 2020 and sending a manned mission to Mars as early as 2030. In response, NASA commissioned the Exploration Systems Architecture Study (ESAS), presenting a full lunar ...

Research paper thumbnail of Investigation of Alternate Transportation Architectures for Crewed Mars Missions

57th International Astronautical Congress, 2006

Research paper thumbnail of IAC-06-D 1 . 1 . 01 On-Orbit Propellant Resupply Options for Mars Exploration Architectures

A report detailing recommendations for a transportation architecture and a roadmap for U.S. explo... more A report detailing recommendations for a transportation architecture and a roadmap for U.S. exploration of the Moon and Mars was released by the NASA Exploration Systems Architecture Study (ESAS) in November 2005. In addition to defining launch vehicles and various aspects of a lunar exploration architecture, the report also elaborated on the extent of commercial involvement in future NASA activities, such as cargo transportation to the International Space Station. Another potential area of commercial involvement under investigation is the delivery of cryogenic propellants to low-Earth orbit (LEO) to refuel NASA assets as well as commercial assets on orbit. The ability to resupply propellant to various architecture elements on-orbit opens a host of new possibilities with respect to a Mars transportation architecture – first and foremost being the ability to conduct a Martian exploration campaign without the development of expensive propulsion systems such as nuclear thermal propulsi...

Research paper thumbnail of Development of a Lunar Architecture Simulation Environment for Evaluating the Use of Propellant Re-Supply

AIAA Modeling and Simulation Technologies Conference and Exhibit, 2007

The NASA Exploration Systems Architecture Study (ESAS) 1 produced a transportation architecture f... more The NASA Exploration Systems Architecture Study (ESAS) 1 produced a transportation architecture for returning humans to the moon affordably and safely, while using commercial services for tasks such as cargo delivery to low earth orbit (LEO). Another potential utilization of commercial services is the delivery of cryogenic propellants to LEO for use in lunar exploration activities. With in-space propellant re-supply available, there is the potential to increase the payload that can be delivered to the lunar surface, increase lunar mission durations, and enable a wider range of lunar missions. The addition of onorbit propellant re-supply would have far-reaching effects on the entire exploration architecture. Currently 70% of the weight delivered to LEO by the cargo launch vehicle is propellant needed for the TLI burn. This is a considerable burden and significantly limits the design freedom of the architecture. The ability of commercial providers to deliver cryogenic propellants to LEO may provide for lower cost and better performing lunar architecture. A model of this architecture has been developed to measure the performance, cost, reliability, mission success rate, and various other criteria for a lunar architecture built around propellant re-supply. This model will provide insight into how the addition of propellant re-supply will affect each aspect of a lunar campaign and help measure the benefits and costs associated with the development and utilization of this capability. The environment itself has been developed using parametric models of the individual architecture elements to allow for the quick evaluation of different architecture trades. A morphological matrix of the different trades considered is provided in Figure 1. This provides an easy to follow outline of the different trades studies conducted using this simulation environment. In addition to being able to quickly asses different architecture designs the simulation environment must be able to provide accurate results. A validation test case was done using the results from the ESAS. The results of this validation showed that the models could predict the results of this architecture within 2-3%. The analysis also includes multi-criteria decision making analysis so that a ranking of the alternatives can be found for a given weighting scenario. This allows for all of the figures of merit to be included in the decision making process. The results of the simulation will allow a decision maker to evaluate the different architecture options against the baseline design and determine what figures of merit are improved with the inclusion of propellant re-supply and if this new capability provides for a better opportunity to meet the needs of future exploration missions.

Research paper thumbnail of An Approach for Calculating the Cost of Launch Vehicle Reliability

AIAA SPACE 2007 Conference & Exposition, 2007

The goal of this paper is to determine the cost of increasing launch vehicle reliability during c... more The goal of this paper is to determine the cost of increasing launch vehicle reliability during conceptual design. The launch vehicle mission requirements are held constant while various reliability strategies are evaluated for their affects on different performance and cost metrics. Traditional design disciplines, such as trajectory analysis and propulsion are included within the performance analysis while the cost discipline focuses on launch vehicle development and production cost. The reliability modeling is developed specifically for application to launch vehicles. A design environment is created that integrates the performance, cost, and reliability disciplines for use with optimization. The integrated environment is utilized to determine a set of optimal design configurations based on a specific weighting of cost and reliability. Different design options for the Cargo Launch Vehicle from the Exploration System Architecture Study are considered and the final result is a set of configurations optimized for a particular weighting of cost and reliability.

Research paper thumbnail of Technology Engineering: The Concurrent Development of Space Transportation Systems and Technology

Engineering Management Journal, 2010

ABSTRACT

Research paper thumbnail of Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference, 2006

Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion... more Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion concepts such as rocket based combined cycle engine (RBCC) and high energy density material (HEDM) propellants. These advanced propulsion elements make the Lazarus launch vehicle both feasible and viable in today's highly competitive market. The Lazarus concept is powered by six rocket based combined cycle engines. These engines are designed to operate with HEDM fuel and liquid oxygen (LOX). During atmospheric flight the LOX is augmented by air traveling through the engines and the resulting propellant mass fractions make single stage to orbit (SSTO) possible. A typical hindrance to SSTO vehicles are the large wings and landing gear necessary for takeoff of a fully fueled vehicle. The Lazarus concept addresses this problem by using a sled to take off horizontally. This sled accelerates the vehicle to over 500 mph using the launch vehicle engines and a propellant cross feed system. This propellant feed system allows the vehicle to accelerate using its own propulsion system without carrying the necessary fuel required while it is attached to the sled. Lazarus is designed to deliver 5,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). This mission design allows for rapid redeployment of small orbital assets with little launch preparation. Lazarus is also designed for a secondary strike mission. The high speed and long range inherent in a SSTO launch vehicle make it an ideal global strike platform. Details of the conceptual design process used for Lazarus are included in this paper. The disciplines used in the design include aerodynamics, configuration, propulsion design, trajectory, mass properties, cost, operations, reliability and safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined into an integrated design process and used to minimize the gross weight of the Lazarus design. Nomenclature α = angle-of-attack, ° AFRSI = Advanced Flexible Reusable Surface Insulation CAD = computer aided design CER = cost estimating relationship c L = coefficient of lift DDT&E = design, development, test, & evaluation DoD = Department of Defense DSM = Design Structure Matrix EMA = electro-mechanical actuators

Research paper thumbnail of A Reduced Order Lunar Lander Model for Rapid Architecture Analysis

AIAA Modeling and Simulation Technologies Conference and Exhibit, 2007

The conceptual design of an architecture for space exploration involves the evaluation of many co... more The conceptual design of an architecture for space exploration involves the evaluation of many concepts. These design spaces may encompass millions or billions of options when each trade is evaluated at the system, vehicle, subsystem, and component level. Various techniques are typically employed to select the configuration of systems that best meets the requirements of the architecture. These include multi-attribute decision making techniques as well as optimization with the use of genetic algorithms and other stochastic methods. In order to speed up the evaluation of these options, a set of reduced-order vehicle models can be used. These models evaluate the gross weight, dry weight, cost, and reliability of a vehicle given a set of programmatic and performance options in less than a second, versus the use of design codes that take on the order of minutes to hours to converge to a vehicle design. The use of such reduced-order models also enables other techniques that would otherwise take too long to run, such as Monte Carlo simulation to model uncertainty, as well as optimization of the vehicle and studies of sensitivities to changes in programmatic and performance inputs. A reduced-order lunar lander model is presented, utilizing response surface equations (RSEs) in place of detailed disciplinary simulations. While some fidelity is lost in approximating these disciplines with RSEs, this approach can be used to evaluate the relative impact of various trade studies at the subsystem, vehicle, and architecture levels. The propulsion system is modeled using a response surface of the REDTOP-2 code. In a similar manner, the trajectory for lunar descent and ascent is simulated using Program to Optimize Simulated Trajectories (POST), and then approximated with a RSE for use in the reduced-order lunar lander model. The weights and sizing model of the lunar lander is based on a combination of historical mass estimating relationships (MERs), and physicsbased mass estimating relationships. Development and production cost modeling is performed using the Cost Estimating Relationships (CERs) from the NASA-Air Force Cost Model (NAFCOM).

Research paper thumbnail of Lunar Module Descent Mission Design

AIAA/AAS Astrodynamics Specialist Conference and Exhibit, 2008

Various lunar descent trajectories were analyzed that include the optimization of the Apollo cons... more Various lunar descent trajectories were analyzed that include the optimization of the Apollo constrained mission trajectory, a fully optimized minimum energy trajectory, and a optimal, constrained trajectory using current instrumentation technology. Trade studies were conducted to determine the impacts of mission assumptions, pilot in the loop/automated flight demands, and additional constraints for the present recurring missions to the same outpost landing site. For mission design at this conceptual phase of the program, the Apollo pre-mission planning was applied to account for known contingencies (hardware, instrumentation known uncertainties) and unknown unknowns. The mission Delta-V's are presented in a risk form of conservative, nominal, and optimistic range where 90 percent of Delta-V is derived by trajectory analysis and the other 10 percent was derived from a qualitative analysis from Apollo 11 pre-mission planning. The recommendations for the Delta Vs are the following: conservative (Apollo derived) (2262 m/s), nominal (2053 m/s), and optimistic (1799 m/s). Because of the qualitative nature of the results, the degree of autonomy assumed, additional safety considerations for a lunar outpost, and the impact of advanced instrumentation, more in-depth analyses are required to refine the present recommendations.

Research paper thumbnail of Plan B for U.S. Human Space Exploration Program

AIAA SPACE 2014 Conference and Exposition, 2014

The Augustine Report recommended the use of commercial launch and propellant depots to improve Pr... more The Augustine Report recommended the use of commercial launch and propellant depots to improve Program sustainability. Plan A Apollo and Apollo programs provide a basis for sustainability where Apollo was quickly cancelled after the lunar landing but the Shuttle was sustained for 40 years. For an asteroid and lunar missions, commercial launch vehicles with depots showed significant cost improvements that are below current NASA exploration budgets with improvement of mission reliability over NASA baseline architectures and missions.

Research paper thumbnail of Understanding the Lunar System Architecture Design Space

AIAA SPACE 2013 Conference and Exposition, 2013

Based on the flexible path strategy and the desire of the international community, the lunar surf... more Based on the flexible path strategy and the desire of the international community, the lunar surface remains a destination for future human exploration. This paper explores options within the lunar system architecture design space, identifying performance requirements placed on the propulsive system that performs Earth departure within that architecture based on existing and/or near-term capabilities. The lander crew module and ascent stage propellant mass fraction are primary drivers for feasibility in multiple lander configurations. As the aggregation location moves further out of the lunar gravity well, the lunar lander is required to perform larger burns, increasing the sensitivity to these two factors. Adding an orbit transfer stage to a two-stage lunar lander and using a large storable stage for braking with a one-stage lunar lander enable higher aggregation locations than Low Lunar Orbit. Finally, while using larger vehicles enables a larger feasible design space, there are still feasible scenarios that use three launches of smaller vehicles.

Research paper thumbnail of Launch Vehicle Propulsion Design with Multiple Selection Criteria

Journal of Spacecraft and Rockets, 2006

Research paper thumbnail of Apollo Lunar Orbit Rendezvous Architecture Decision Revisited

Journal of Spacecraft and Rockets, 2006

The 1962 Apollo architecture mode decision process was revisited with modern analysis and systems... more The 1962 Apollo architecture mode decision process was revisited with modern analysis and systems engineer tools to determine driving selection criteria and technology/operational mode design decisions that may be used for NASA's current Space Exploration program. Results of the study agreed with the Apollo selection of the Lunar Orbit Rendezvous mode based on the technology maturity and politics in 1962. Using today's greater emphasis on human safety and improvements in technology and design maturity, a slight edge may be given to the direct lunar mode over lunar orbit rendezvous. Also, the NOVA direct mode and Earth orbit rendezvous mode are not competitive based any selection criteria. Finally, reliability and development, operations, and production costs are major drivers in today's decision process.

Research paper thumbnail of Launch Vehicle Engine Selection Using Probabilistic Techniques

48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 2007

A new method for selecting the number of engines on a rocket stage based upon reliability and cos... more A new method for selecting the number of engines on a rocket stage based upon reliability and cost is presented. This method will compare a new technique for reliability analysis that results in a higher fidelity model when considering engine out capability. The cost of each vehicle configuration is calculated to find an optimal balance of system cost and reliability. The optimal solution will be determined by using a combined objective function based on minimizing campaign cost and maximizing vehicle reliability. When performing reliability analysis with conventional practices, an engine out scenario is calculated using static tools. Another technique presented here will use a model that can adjust the failure rate of the propulsion subsystem to account for a longer burn time when an engine fails. The propulsion reliability will be combined with the other subsystem reliability estimates to calculate the reliability of the overall system. The system reliability will be one part of an objective function used in the optimization scheme. The methodology is created by using a combination of industry standard tools along with author developed models to create an integrated framework that allows for optimization. The complete design space is explored using optimization to examine the maximum reliability and minimum cost configurations. Additionally, Monte Carlo Simulation is used to vary the single engine failure rate. The ranges for the simulation are drawn from historical data and are used to capture the effects of a higher fidelity modeling technique. Conclusions are made about the engine configuration on the S-II stage of the Saturn V based upon the methodology presented in this paper.

Research paper thumbnail of A Method for Allocating Reliability and Cost in a Lunar Exploration Architecture

In January 2005, President Bush announced the Vision for Space Exploration. This vision involved ... more In January 2005, President Bush announced the Vision for Space Exploration. This vision involved a progressive expansion of human capabilities beyond Low Earth Orbit. Current design processes utilized to meet this vision employ performance based optimization schemes to determine the ideal solution. In these design processes the important aspects such as cost and reliability are currently calculated as an afterthought to the traditional performance metrics. The methodology implemented in this paper focuses on bringing these decisive variables to the forefront of the design process. To achieve this focus on cost and reliability in a lunar architecture design, a resource allocation technique from the business world will be implemented. This allocation technique optimally distributes the company’s resources even though the actual performances of the products are uncertain. This method of resource allocation will be applied to a lunar architecture design to achieve the highest architectu...

Research paper thumbnail of Launch System Trajectory Performance Modeling and Sensitivity Analysis Using Response Surface Methods

Research paper thumbnail of Spare Strategy Analysis for Life Support Systems for Human Space Exploration

Journal of Spacecraft and Rockets

Research paper thumbnail of Dual-fuel propulsion - Why it works, possible engines, and results of vehicle studies149>on earth-to-orbit Space Shuttle flights

Conference on Advanced Technology for Future Space Systems

ABSTRACT

Research paper thumbnail of Effects of Relaxed Static Longitudinal Stability on a Single-Stage-to-Orbit Vehicle Design

... Descriptive Note : Technical paper,. Corporate Author : NATIONAL AERONAUTICS AND SPACE ADMINI... more ... Descriptive Note : Technical paper,. Corporate Author : NATIONAL AERONAUTICS AND SPACE ADMINISTRATION HAMPTON VA LANGLEY RESEARCH CENTER. Personal Author(s) : Freeman,Delma C. , Jr. ; Wilhite,Alan W. Report Date : DEC 1979. ...

Research paper thumbnail of Design options for advanced manned launch systems (AMLS)

Various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to ... more Various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to Space Station and polar orbit. Included are single- and two-stage winged systems with rocket and/or airbreathing propulsion systems. For near-term technologies, two-stage, reusable rocket systems are favored over single-stage rocket or two-stage airbreathing/rocket systems. Advanced technologies enable viable single-stage-to-orbit (SSTO) concepts. Although two-stage rocket systems continue

Research paper thumbnail of ESAS-Derived Earth Departure Stage Design for Human Mars Exploration

Retrieved …, 2007

Background The Vision for Space Exploration mandated the goal of returning humans to the Moon by ... more Background The Vision for Space Exploration mandated the goal of returning humans to the Moon by 2020 and sending a manned mission to Mars as early as 2030. In response, NASA commissioned the Exploration Systems Architecture Study (ESAS), presenting a full lunar ...

Research paper thumbnail of Investigation of Alternate Transportation Architectures for Crewed Mars Missions

57th International Astronautical Congress, 2006

Research paper thumbnail of IAC-06-D 1 . 1 . 01 On-Orbit Propellant Resupply Options for Mars Exploration Architectures

A report detailing recommendations for a transportation architecture and a roadmap for U.S. explo... more A report detailing recommendations for a transportation architecture and a roadmap for U.S. exploration of the Moon and Mars was released by the NASA Exploration Systems Architecture Study (ESAS) in November 2005. In addition to defining launch vehicles and various aspects of a lunar exploration architecture, the report also elaborated on the extent of commercial involvement in future NASA activities, such as cargo transportation to the International Space Station. Another potential area of commercial involvement under investigation is the delivery of cryogenic propellants to low-Earth orbit (LEO) to refuel NASA assets as well as commercial assets on orbit. The ability to resupply propellant to various architecture elements on-orbit opens a host of new possibilities with respect to a Mars transportation architecture – first and foremost being the ability to conduct a Martian exploration campaign without the development of expensive propulsion systems such as nuclear thermal propulsi...

Research paper thumbnail of Development of a Lunar Architecture Simulation Environment for Evaluating the Use of Propellant Re-Supply

AIAA Modeling and Simulation Technologies Conference and Exhibit, 2007

The NASA Exploration Systems Architecture Study (ESAS) 1 produced a transportation architecture f... more The NASA Exploration Systems Architecture Study (ESAS) 1 produced a transportation architecture for returning humans to the moon affordably and safely, while using commercial services for tasks such as cargo delivery to low earth orbit (LEO). Another potential utilization of commercial services is the delivery of cryogenic propellants to LEO for use in lunar exploration activities. With in-space propellant re-supply available, there is the potential to increase the payload that can be delivered to the lunar surface, increase lunar mission durations, and enable a wider range of lunar missions. The addition of onorbit propellant re-supply would have far-reaching effects on the entire exploration architecture. Currently 70% of the weight delivered to LEO by the cargo launch vehicle is propellant needed for the TLI burn. This is a considerable burden and significantly limits the design freedom of the architecture. The ability of commercial providers to deliver cryogenic propellants to LEO may provide for lower cost and better performing lunar architecture. A model of this architecture has been developed to measure the performance, cost, reliability, mission success rate, and various other criteria for a lunar architecture built around propellant re-supply. This model will provide insight into how the addition of propellant re-supply will affect each aspect of a lunar campaign and help measure the benefits and costs associated with the development and utilization of this capability. The environment itself has been developed using parametric models of the individual architecture elements to allow for the quick evaluation of different architecture trades. A morphological matrix of the different trades considered is provided in Figure 1. This provides an easy to follow outline of the different trades studies conducted using this simulation environment. In addition to being able to quickly asses different architecture designs the simulation environment must be able to provide accurate results. A validation test case was done using the results from the ESAS. The results of this validation showed that the models could predict the results of this architecture within 2-3%. The analysis also includes multi-criteria decision making analysis so that a ranking of the alternatives can be found for a given weighting scenario. This allows for all of the figures of merit to be included in the decision making process. The results of the simulation will allow a decision maker to evaluate the different architecture options against the baseline design and determine what figures of merit are improved with the inclusion of propellant re-supply and if this new capability provides for a better opportunity to meet the needs of future exploration missions.

Research paper thumbnail of An Approach for Calculating the Cost of Launch Vehicle Reliability

AIAA SPACE 2007 Conference & Exposition, 2007

The goal of this paper is to determine the cost of increasing launch vehicle reliability during c... more The goal of this paper is to determine the cost of increasing launch vehicle reliability during conceptual design. The launch vehicle mission requirements are held constant while various reliability strategies are evaluated for their affects on different performance and cost metrics. Traditional design disciplines, such as trajectory analysis and propulsion are included within the performance analysis while the cost discipline focuses on launch vehicle development and production cost. The reliability modeling is developed specifically for application to launch vehicles. A design environment is created that integrates the performance, cost, and reliability disciplines for use with optimization. The integrated environment is utilized to determine a set of optimal design configurations based on a specific weighting of cost and reliability. Different design options for the Cargo Launch Vehicle from the Exploration System Architecture Study are considered and the final result is a set of configurations optimized for a particular weighting of cost and reliability.

Research paper thumbnail of Technology Engineering: The Concurrent Development of Space Transportation Systems and Technology

Engineering Management Journal, 2010

ABSTRACT

Research paper thumbnail of Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference, 2006

Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion... more Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion concepts such as rocket based combined cycle engine (RBCC) and high energy density material (HEDM) propellants. These advanced propulsion elements make the Lazarus launch vehicle both feasible and viable in today's highly competitive market. The Lazarus concept is powered by six rocket based combined cycle engines. These engines are designed to operate with HEDM fuel and liquid oxygen (LOX). During atmospheric flight the LOX is augmented by air traveling through the engines and the resulting propellant mass fractions make single stage to orbit (SSTO) possible. A typical hindrance to SSTO vehicles are the large wings and landing gear necessary for takeoff of a fully fueled vehicle. The Lazarus concept addresses this problem by using a sled to take off horizontally. This sled accelerates the vehicle to over 500 mph using the launch vehicle engines and a propellant cross feed system. This propellant feed system allows the vehicle to accelerate using its own propulsion system without carrying the necessary fuel required while it is attached to the sled. Lazarus is designed to deliver 5,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). This mission design allows for rapid redeployment of small orbital assets with little launch preparation. Lazarus is also designed for a secondary strike mission. The high speed and long range inherent in a SSTO launch vehicle make it an ideal global strike platform. Details of the conceptual design process used for Lazarus are included in this paper. The disciplines used in the design include aerodynamics, configuration, propulsion design, trajectory, mass properties, cost, operations, reliability and safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined into an integrated design process and used to minimize the gross weight of the Lazarus design. Nomenclature α = angle-of-attack, ° AFRSI = Advanced Flexible Reusable Surface Insulation CAD = computer aided design CER = cost estimating relationship c L = coefficient of lift DDT&E = design, development, test, & evaluation DoD = Department of Defense DSM = Design Structure Matrix EMA = electro-mechanical actuators

Research paper thumbnail of A Reduced Order Lunar Lander Model for Rapid Architecture Analysis

AIAA Modeling and Simulation Technologies Conference and Exhibit, 2007

The conceptual design of an architecture for space exploration involves the evaluation of many co... more The conceptual design of an architecture for space exploration involves the evaluation of many concepts. These design spaces may encompass millions or billions of options when each trade is evaluated at the system, vehicle, subsystem, and component level. Various techniques are typically employed to select the configuration of systems that best meets the requirements of the architecture. These include multi-attribute decision making techniques as well as optimization with the use of genetic algorithms and other stochastic methods. In order to speed up the evaluation of these options, a set of reduced-order vehicle models can be used. These models evaluate the gross weight, dry weight, cost, and reliability of a vehicle given a set of programmatic and performance options in less than a second, versus the use of design codes that take on the order of minutes to hours to converge to a vehicle design. The use of such reduced-order models also enables other techniques that would otherwise take too long to run, such as Monte Carlo simulation to model uncertainty, as well as optimization of the vehicle and studies of sensitivities to changes in programmatic and performance inputs. A reduced-order lunar lander model is presented, utilizing response surface equations (RSEs) in place of detailed disciplinary simulations. While some fidelity is lost in approximating these disciplines with RSEs, this approach can be used to evaluate the relative impact of various trade studies at the subsystem, vehicle, and architecture levels. The propulsion system is modeled using a response surface of the REDTOP-2 code. In a similar manner, the trajectory for lunar descent and ascent is simulated using Program to Optimize Simulated Trajectories (POST), and then approximated with a RSE for use in the reduced-order lunar lander model. The weights and sizing model of the lunar lander is based on a combination of historical mass estimating relationships (MERs), and physicsbased mass estimating relationships. Development and production cost modeling is performed using the Cost Estimating Relationships (CERs) from the NASA-Air Force Cost Model (NAFCOM).

Research paper thumbnail of Lunar Module Descent Mission Design

AIAA/AAS Astrodynamics Specialist Conference and Exhibit, 2008

Various lunar descent trajectories were analyzed that include the optimization of the Apollo cons... more Various lunar descent trajectories were analyzed that include the optimization of the Apollo constrained mission trajectory, a fully optimized minimum energy trajectory, and a optimal, constrained trajectory using current instrumentation technology. Trade studies were conducted to determine the impacts of mission assumptions, pilot in the loop/automated flight demands, and additional constraints for the present recurring missions to the same outpost landing site. For mission design at this conceptual phase of the program, the Apollo pre-mission planning was applied to account for known contingencies (hardware, instrumentation known uncertainties) and unknown unknowns. The mission Delta-V's are presented in a risk form of conservative, nominal, and optimistic range where 90 percent of Delta-V is derived by trajectory analysis and the other 10 percent was derived from a qualitative analysis from Apollo 11 pre-mission planning. The recommendations for the Delta Vs are the following: conservative (Apollo derived) (2262 m/s), nominal (2053 m/s), and optimistic (1799 m/s). Because of the qualitative nature of the results, the degree of autonomy assumed, additional safety considerations for a lunar outpost, and the impact of advanced instrumentation, more in-depth analyses are required to refine the present recommendations.

Research paper thumbnail of Plan B for U.S. Human Space Exploration Program

AIAA SPACE 2014 Conference and Exposition, 2014

The Augustine Report recommended the use of commercial launch and propellant depots to improve Pr... more The Augustine Report recommended the use of commercial launch and propellant depots to improve Program sustainability. Plan A Apollo and Apollo programs provide a basis for sustainability where Apollo was quickly cancelled after the lunar landing but the Shuttle was sustained for 40 years. For an asteroid and lunar missions, commercial launch vehicles with depots showed significant cost improvements that are below current NASA exploration budgets with improvement of mission reliability over NASA baseline architectures and missions.

Research paper thumbnail of Understanding the Lunar System Architecture Design Space

AIAA SPACE 2013 Conference and Exposition, 2013

Based on the flexible path strategy and the desire of the international community, the lunar surf... more Based on the flexible path strategy and the desire of the international community, the lunar surface remains a destination for future human exploration. This paper explores options within the lunar system architecture design space, identifying performance requirements placed on the propulsive system that performs Earth departure within that architecture based on existing and/or near-term capabilities. The lander crew module and ascent stage propellant mass fraction are primary drivers for feasibility in multiple lander configurations. As the aggregation location moves further out of the lunar gravity well, the lunar lander is required to perform larger burns, increasing the sensitivity to these two factors. Adding an orbit transfer stage to a two-stage lunar lander and using a large storable stage for braking with a one-stage lunar lander enable higher aggregation locations than Low Lunar Orbit. Finally, while using larger vehicles enables a larger feasible design space, there are still feasible scenarios that use three launches of smaller vehicles.

Research paper thumbnail of Launch Vehicle Propulsion Design with Multiple Selection Criteria

Journal of Spacecraft and Rockets, 2006

Research paper thumbnail of Apollo Lunar Orbit Rendezvous Architecture Decision Revisited

Journal of Spacecraft and Rockets, 2006

The 1962 Apollo architecture mode decision process was revisited with modern analysis and systems... more The 1962 Apollo architecture mode decision process was revisited with modern analysis and systems engineer tools to determine driving selection criteria and technology/operational mode design decisions that may be used for NASA's current Space Exploration program. Results of the study agreed with the Apollo selection of the Lunar Orbit Rendezvous mode based on the technology maturity and politics in 1962. Using today's greater emphasis on human safety and improvements in technology and design maturity, a slight edge may be given to the direct lunar mode over lunar orbit rendezvous. Also, the NOVA direct mode and Earth orbit rendezvous mode are not competitive based any selection criteria. Finally, reliability and development, operations, and production costs are major drivers in today's decision process.

Research paper thumbnail of Launch Vehicle Engine Selection Using Probabilistic Techniques

48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 2007

A new method for selecting the number of engines on a rocket stage based upon reliability and cos... more A new method for selecting the number of engines on a rocket stage based upon reliability and cost is presented. This method will compare a new technique for reliability analysis that results in a higher fidelity model when considering engine out capability. The cost of each vehicle configuration is calculated to find an optimal balance of system cost and reliability. The optimal solution will be determined by using a combined objective function based on minimizing campaign cost and maximizing vehicle reliability. When performing reliability analysis with conventional practices, an engine out scenario is calculated using static tools. Another technique presented here will use a model that can adjust the failure rate of the propulsion subsystem to account for a longer burn time when an engine fails. The propulsion reliability will be combined with the other subsystem reliability estimates to calculate the reliability of the overall system. The system reliability will be one part of an objective function used in the optimization scheme. The methodology is created by using a combination of industry standard tools along with author developed models to create an integrated framework that allows for optimization. The complete design space is explored using optimization to examine the maximum reliability and minimum cost configurations. Additionally, Monte Carlo Simulation is used to vary the single engine failure rate. The ranges for the simulation are drawn from historical data and are used to capture the effects of a higher fidelity modeling technique. Conclusions are made about the engine configuration on the S-II stage of the Saturn V based upon the methodology presented in this paper.

Research paper thumbnail of A Method for Allocating Reliability and Cost in a Lunar Exploration Architecture

In January 2005, President Bush announced the Vision for Space Exploration. This vision involved ... more In January 2005, President Bush announced the Vision for Space Exploration. This vision involved a progressive expansion of human capabilities beyond Low Earth Orbit. Current design processes utilized to meet this vision employ performance based optimization schemes to determine the ideal solution. In these design processes the important aspects such as cost and reliability are currently calculated as an afterthought to the traditional performance metrics. The methodology implemented in this paper focuses on bringing these decisive variables to the forefront of the design process. To achieve this focus on cost and reliability in a lunar architecture design, a resource allocation technique from the business world will be implemented. This allocation technique optimally distributes the company’s resources even though the actual performances of the products are uncertain. This method of resource allocation will be applied to a lunar architecture design to achieve the highest architectu...

Research paper thumbnail of Launch System Trajectory Performance Modeling and Sensitivity Analysis Using Response Surface Methods

Research paper thumbnail of Spare Strategy Analysis for Life Support Systems for Human Space Exploration

Journal of Spacecraft and Rockets

Research paper thumbnail of Dual-fuel propulsion - Why it works, possible engines, and results of vehicle studies149>on earth-to-orbit Space Shuttle flights

Conference on Advanced Technology for Future Space Systems

ABSTRACT

Research paper thumbnail of Effects of Relaxed Static Longitudinal Stability on a Single-Stage-to-Orbit Vehicle Design

... Descriptive Note : Technical paper,. Corporate Author : NATIONAL AERONAUTICS AND SPACE ADMINI... more ... Descriptive Note : Technical paper,. Corporate Author : NATIONAL AERONAUTICS AND SPACE ADMINISTRATION HAMPTON VA LANGLEY RESEARCH CENTER. Personal Author(s) : Freeman,Delma C. , Jr. ; Wilhite,Alan W. Report Date : DEC 1979. ...

Research paper thumbnail of Design options for advanced manned launch systems (AMLS)

Various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to ... more Various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to Space Station and polar orbit. Included are single- and two-stage winged systems with rocket and/or airbreathing propulsion systems. For near-term technologies, two-stage, reusable rocket systems are favored over single-stage rocket or two-stage airbreathing/rocket systems. Advanced technologies enable viable single-stage-to-orbit (SSTO) concepts. Although two-stage rocket systems continue