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Aborts during the critical ascent flight phase require the design and operation of Orion Crew Exploration Vehicle (CEV) systems to escape from the Crew Launch Vehicle (CLV) and return the crew safely to the Earth. To accomplish this requirement of continuous abort coverage, CEV ascent abort modes are being designed and analyzed to accommodate the velocity, altitude, atmospheric, and vehicle configuration changes that occur during ascent. Aborts from the launch pad to early in the flight of the CLV second stage are performed using the Launch Abort System (LAS). During this type of abort, the LAS Abort Motor is used to pull the Crew Module (CM) safely away from the CLV and Service Module (SM). LAS abort guidance and control studies and design trades are being conducted so that more informed decisions can be made regarding the vehicle abort requirements, design, and operation. This paper presents an overview of the Orion CEV, an overview of the LAS ascent abort mode, and a summary of key LAS abort analysis methods and results.

A multiple spacecraft close-proximity control algorithm was implemented and tested with the Synchronized Position Hold Engage and Reorient Experimental Satellites (SPHERES) facility onboard the International Space Station (ISS). During flight testing, a chaser satellite successfully approached a virtual target satellite, while avoiding collision with a virtual obstacle satellite. This research contributes to the control of multiple spacecraft for emerging missions, which may require simultaneous gathering, rendezvous, and docking. The unique control algorithm was developed at NPS and integrated onto the MIT SPHERES facility. The control algorithm implemented combines the efficiency of the Linear Quadratic Regulator (LQR), and the robust collision avoidance capability of the Artificial Potential Function method (APF). The LQR control effort serves as the attractive force toward goal positions, while the APF-based repulsive functions provide collision avoidance for both fixed and moving obstacles. The amalgamation of these two control methods into a multiple spacecraft close-proximity control algorithm yielded promising results as demonstrated by simulations performed at NPS. Comprehensive simulation evaluation enabled implementation and testing of the spacecraft control algorithm on the SPHERES facility at MIT. Finally, successful ground testing enabled execution of flight testing onboard the ISS. The NPS’s Spacecraft Robotics Laboratory (SRL) and MIT’s Space Systems Laboratory (SSL) simulations, the MIT’s SSL SPHERES ground testing, and the SPHERES flight testing results are all presented in this paper.

The Orion Capsule is designed for two types of missions: 1) Low Earth Orbit missions to the ISS, and 2) missions to the moon and perhaps beyond. The constraints on the vehicle during atmospheric re-entry from either of these missions include constraints on heating rate, total heat load, maximum gload limits, gload duration limits, and landing site accuracy requirements associated with nominal and back-up mode entries. Also to be considered is the proper disposal of the Service Module for both ISS and lunar missions, as well as a docking mechanism for ISS return. For ISS class missions, the capsule will use an Apollo-style direct entry. For lunar return missions, depending on the timing of the mission, the capsule could perform a direct entry or a skip entry of up to 4800 n.mi. in order to land in the coastal waters of California. The mission requirements and vehicle constraints for the various mission scenarios must all be considered when designing an entry ∞ight corridor. A standardized process for creating and analyzing the entry ∞ight corridor for all possible mission scenarios has been created to aid in the design of the Orion vehicle. The process involves a series of Monte Carlo simulations for the nominal and back-up modes for a particular entry scenario. The results from these simulations are then analyzed to determine acceptability for constraint margins and landing site accuracy. Several Monte Carlo suites are performed to flnd the acceptable boundaries of the major entry ∞ight corridor parameters, including the initial ∞ight path angle, initial pre-bank angle, crossrange targeting, and entry guidance parameters.

The analysis of candidate thruster configurations for the Crew Exploration Vehicle (CEV) is presented. Six candidate configurations were considered for the prime contractor baseline design. The analysis included analytical assessments of control authority, control precision, efficiency and robustness, as well as simulation assessments of control performance. The principles used in the analytic assessments of controllability, robustness and fuel performance are covered and results provided for the configurations assessed. Simulation analysis was conducted using a pulse width modulated, 6 DOF reaction system control law with a simplex-based thruster selection algorithm. Control laws were automatically derived from hardware configuration parameters including thruster locations, directions, magnitude and specific impulse, as well as vehicle mass properties. This parameterized controller allowed rapid assessment of multiple candidate layouts. Simulation results are presented for final phase rendezvous and docking, as well as low lunar orbit attitude hold. Finally, on-going analysis to consider alternate Service Module designs and to assess the pilot-ability of the baseline design are discussed to provide a status of orbit control design work to date.

The Orion Spacecraft will be required to perform entry and landing functions for both Low Earth Orbit (LEO) and Lunar return missions, utilizing only the Command Module (CM) with its unique systems and GN&C design. This paper presents the current CM Flight Control System (FCS) design to support entry and landing, with a focus on analyses that have supported its development to date. The CM FCS will have to provide for spacecraft stability and control while following guidance or manual commands during exo-atmospheric flight, after Service Module separation, translational powered flight required of the CM, atmospheric flight supporting both direct entry and skip trajectories down to drogue chute deploy, and during roll attitude reorientation just prior to touchdown. Various studies and analyses have been performed or are on-going supporting an overall FCS design with reasonably sized Reaction Control System (RCS) jets, that minimizes fuel usage, that provides appropriate command following but with reasonable stability and control margin. Results from these efforts to date are included, with particular attention on design issues that have emerged, such as the struggle to accommodate sub-sonic pitch and yaw control without using excessively large jets that could have a detrimental impact on vehicle weight. Apollo, with a similar shape, struggled with this issue as well. Outstanding CM FCS related design and analysis issues, planned for future effort, are also briefly be discussed.

The Orion Crew Exploration Vehicle is the first crewed capsule design to be developed by NASA since Project Apollo. Unlike Apollo, however, the CEV is being designed for service in both Lunar and International Space Station missions. Ascent aborts pose some issues that were not present for Apollo, due to its launch azimuth, nor Space Shuttle, due to its cross range capability. The requirement that a North Atlantic splashdown following an abort be avoidable, in conjunction with the requirement for overlapping abort modes to maximize crew survivability, drives the thrust level of the service module main engine. This paper summarizes 3DOF analysis conducted by NASA to aid in the determination of the appropriate propulsion system for the service module, and the appropriate propellant loading for ISS missions such that crew survivability is maximized.

The Crew Exploration Vehicle (CEV) is required to maintain continuous abort capability from lift off through destination arrival. This requirement is driven by the desire to provide the capability to safely return the crew to Earth after failure scenarios during the various phases of the mission. This paper addresses abort trajectory design considerations, concept of operations and guidance algorithm prototypes for the portion of the ascent trajectory following nominal jettison of the Launch Abort System (LAS) until safe orbit insertion. Factors such as abort system performance, crew load limits, natural environments, crew recovery, and vehicle element disposal were investigated to determine how to achieve continuous vehicle abort capability.