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A general overview of the Orion abort flight test is presented. The contents include: 1) Abort Flight Test Project Overview; 2) DFRC Exploration Mission Directorate; 3) Abort Flight Test; 4) Flight Test Configurations; 5) Flight Test Vehicle Engineering Office; 6) DFRC FTA Scope; 7) Flight Test Operations; 8) DFRC Ops Support; 9) Launch Facilities; and 10) Scope of Launch Abort Flight Test
This paper describes initial flight performance analyses conducted early in the Orion Project to support concept feasibility studies for the Crew Exploration Vehicle s Launch Abort System (LAS). Key performance requirements that significantly affect abort capability are presented. These requirements have implications on sizing the Abort Motor, tailoring its thrust profile to meet escape requirements for both launch pad and high drag/high dynamic pressure ascent aborts. Additional performance considerations are provided for the Attitude Control Motor, a key element of the Orion LAS design that eliminates the need for ballast and provides performance robustness over a passive control approach. Finally, performance of the LAS jettison function is discussed, along with implications on Jettison Motor sizing and the timing of the jettison event during a nominal mission. These studies provide an initial understanding of LAS performance that will continue to evolve as the Orion design is matured.
The Pad Abort One flight test of the Orion Abort Flight Test Program is currently under development with the goal of demonstrating the capability of the Launch Abort System. In the event of a launch failure, this system will propel the Crew Exploration Vehicle to safety. An essential component of this flight test is range safety, which ensures the security of range assets and personnel. A debris catalog analysis was done as part of a range safety data package delivered to the White Sands Missile Range in New Mexico where the test will be conducted. The analysis discusses the consequences of an overpressurization of the Abort Motor. The resulting structural failure was assumed to create a debris field of vehicle fragments that could potentially pose a hazard to the range. A statistical model was used to assemble the debris catalog of potential propellant fragments. Then, a thermodynamic, energy balance model was applied to the system in order to determine the imparted velocity to these propellant fragments. This analysis was conducted at four points along the flight trajectory to better understand the failure consequences over the entire flight. The methods used to perform this analysis are outlined in detail and the corresponding results are presented and discussed.
The first flight test of the Orion Abort Flight Test project is scheduled to launch in Spring 2010. This flight test is known as Pad Abort 1 (PA-1) and it is intended to accomplish a series of flight test objectives, including demonstrating the capability of the Launch Abort System (LAS) to propel the Crew Module (CM) to a safe distance from a launch vehicle during a pad abort. The PA-1 Flight Test Article (FTA) is actively controlled by a guidance, navigation, and control (GN&C) system for much of its flight. The purpose of this paper is to describe the design, development, and analysis of the PA-1 GN&C system. A description of the technical solutions that were developed to meet the challenge of satisfying many competing requirements is presented. A historical perspective of how the Orion LAV compares to the Apollo Launch Escape Vehicle (LEV) design will also be included.
In order to support the anytime abort requirements of a manned mission to the Moon, the vehicle abort capabilities for the translunar and circumlunar phases of the mission must be studied. Depending on the location of the abort maneuver, the maximum return time to Earth and the available propellant, two different kinds of return trajectories can be calculated: direct and fly-by. This paper presents a new method to compute these return trajectories in a deterministic and fast way without using numerical optimizers. Since no simplifications of the gravity model are required, the resulting trajectories are very accurate and can be used for both mission design and operations. This technique has been extensively used to evaluate the abort capabilities of the Orion/Altair vehicles in the Constellation program for the translunar phase of the mission.
The National Aeronautics and Space Administration (NASA) Engineering and Safety Center (NESC) designed, developed and flew the alternative Max Launch Abort System (MLAS) as risk mitigation for the baseline Orion spacecraft launch abort system (LAS) already in development. The NESC was tasked with both formulating a conceptual objective system (OS) design of this alternative MLAS as well as demonstrating this concept with a simulated pad abort flight test. The goal was to obtain sufficient flight test data to assess performance, validate models/tools, and to reduce the design and development risks for a MLAS OS. Less than 2 years after Project start the MLAS simulated pad abort flight test was successfully conducted from Wallops Island on July 8, 2009. The entire flight test duration was 88 seconds during which time multiple staging events were performed and nine separate critically timed parachute deployments occurred as scheduled. Overall, the as-flown flight performance was as predicted prior to launch. This paper provides an overview of the guidance navigation and control (GN&C) technical approaches employed on this rapid prototyping activity. This paper describes the methodology used to design the MLAS flight test vehicle (FTV). Lessons that were learned during this rapid prototyping project are also summarized.
The National Aeronautics and Space Administration Engineering and Safety Center designed, developed and flew the alternative Max Launch Abort System (MLAS) as risk mitigation for the baseline Orion spacecraft launch abort system already in development. The NESC was tasked with both formulating a conceptual objective system design of this alternative MLAS as well as demonstrating this concept with a simulated pad abort flight test. Less than 2 years after Project start the MLAS simulated pad abort flight test was successfully conducted from Wallops Island on July 8, 2009. The entire flight test duration was 88 seconds during which time multiple staging events were performed and nine separate critically timed parachute deployments occurred as scheduled. This paper provides an overview of the guidance navigation and control technical approaches employed on this rapid prototyping activity; describes the methodology used to design the MLAS flight test vehicle; and lessons that were learned during this rapid prototyping project are also summarized.
The National Aeronautics and Space Administration (NASA) Engineering and Safety Center (NESC) was tasked to develop an alternate, tower-less launch abort system (LAS) as risk mitigation for the Orion Project. The successful pad abort flight demonstration test in July 2009 of the "Max" launch abort system (MLAS) provided data critical to the design of future LASs, while demonstrating the Agency s ability to rapidly design, build and fly full-scale hardware at minimal cost in a "virtual" work environment. Limited funding and an aggressive schedule presented a challenge for testing of the complex MLAS system. The successful pad abort flight demonstration test was attributed to the project s systems engineering and integration process, which included: a concise definition of, and an adherence to, flight test objectives; a solid operational concept; well defined performance requirements, and a test program tailored to reducing the highest flight test risks. The testing ranged from wind tunnel validation of computational fluid dynamic simulations to component ground tests of the highest risk subsystems. This paper provides an overview of the testing/risk management approach and methodologies used to understand and reduce the areas of highest risk - resulting in a successful flight demonstration test.
The aerodynamic database for the Orion Launch Abort System (LAS) was developed largely from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamics (CFD) simulations. The LAS contains three solid rocket motors used in various phases of an abort to provide propulsion, steering, and Launch Abort Tower (LAT) jettison from the Crew Module (CM). This paper describes a pair of wind tunnel experiments performed at transonic and supersonic speeds to determine the aerodynamic effects due to proximity and jet interactions during LAT jettison from the CM at the end of an abort. The tests were run using two different scale models at angles of attack from 150deg to 200deg , sideslip angles from -10deg to +10deg , and a range of powered thrust levels from the jettison motors to match various jet simulation parameters with flight values. Separation movements between the CM and LAT included axial and vertical translations as well as relative pitch angle between the two bodies. The paper details aspects of the model design, nozzle scaling methodology, instrumentation, testing procedures, and data reduction. Sample data are shown to highlight trends seen in the results.
Modeling the aerodynamics of the Orion Launch Abort Vehicle (LAV) has presented many technical challenges to the developers of the Orion aerodynamic database. During a launch abort event, the aerodynamic environment around the LAV is very complex as multiple solid rocket plumes interact with each other and the vehicle. It is further complicated by vehicle separation events such as between the LAV and the launch vehicle stack or between the launch abort tower and the crew module. The aerodynamic database for the LAV was developed mainly from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamic simulations. However, limitations in both methods have made it difficult to properly capture the aerodynamics of the LAV in experimental and numerical simulations. These limitations have also influenced decisions regarding the modeling and structure of the aerodynamic database for the LAV and led to compromises and creative solutions. Two database modeling approaches are presented in this paper (incremental aerodynamics and total aerodynamics), with examples showing strengths and weaknesses of each approach. In addition, the unique problems presented to the database developers by the large data space required for modeling a launch abort event illustrate the complexities of working with multi-dimensional data.
One of NASAs challenges for the Orion vehicle is the control system design for the Launch Abort Vehicle (LAV), which is required to abort safely at any time during the atmospheric ascent portion of ight. The focus of this paper is the gain design and scheduling process for a controller that covers the wide range of vehicle configurations and flight conditions experienced during the full envelope of potential abort trajectories from the pad to exo-atmospheric flight. Several factors are taken into account in the automation process for tuning the gains including the abort effectors, the environmental changes and the autopilot modes. Gain scheduling is accomplished using a linear quadratic regulator (LQR) approach for the decoupled, simplified linear model throughout the operational envelope in time, altitude and Mach number. The derived gains are then implemented into the full linear model for controller requirement validation. Finally, the gains are tested and evaluated in a non-linear simulation using the vehicles ight software to ensure performance requirements are met. An overview of the LAV controller design and a description of the linear plant models are presented. Examples of the most significant challenges with the automation of the gain tuning process are then discussed. In conclusion, the paper will consider the lessons learned through out the process, especially in regards to automation, and examine the usefulness of the gain scheduling tool and process developed as applicable to non-Orion vehicles.
Wind tunnel tests were conducted by Nielsen Engineering & Research (NEAR) and Rose Engineering & Research (REAR) in conjunction with the NASA Engineering & Safety Center (NESC) on a 6%-scale model of the Orion launch abort vehicle (LAV) configured with four grid fins mounted near the base of the vehicle. The objectives of these tests were to 1) quantify LAV stability augmentation provided by the grid fins from subsonic through supersonic Mach numbers, 2) assess the benefits of swept grid fins versus unswept grid fins on the LAV, 3) determine the effects of the LAV abort motors on grid fin aerodynamics, and 4) generate an aerodynamic database for use in the future application of grid fins to small length-to-diameter ratio vehicles similar to the LAV. The tests were conducted in NASA Ames Research Center's 11x11-foot transonic wind tunnel from Mach 0.5 through Mach 1.3 and in their 9x7-foot supersonic wind tunnel from Mach 1.6 through Mach 2.5. Force- and moment-coefficient data were collected for the complete vehicle and for each individual grid fin as a function of angle of attack and sideslip angle. Tests were conducted with both swept and unswept grid fins with the simulated abort motors (cold jets) off and on. The swept grid fins were designed with a 22.5deg aft sweep angle for both the frame and the internal lattice so that the frontal projection of the swept fins was the same as for the unswept fins. Data from these tests indicate that both unswept and swept grid fins provide significant improvements in pitch stability as compared to the baseline vehicle over the Mach number range investigated. The swept fins typically provide improved stability as compared to the unswept fins, but the performance gap diminished as Mach number was increased. The aerodynamic performance of the fins was not observed to degrade when the abort motors were turned on. Results from these tests indicate that grid fins can be a robust solution for stabilizing the Orion LAV over a wide range of operating conditions.
Launch vehicle ascent is a time of high risk for an on-board crew. There are many types of failures that can kill the crew if the crew is still on-board when the failure becomes catastrophic. For some failure scenarios, there is plenty of time for the crew to be warned and to depart, whereas in some there is insufficient time for the crew to escape. There is a large fraction of possible failures for which time is of the essence and a successful abort is possible if the detection and action happens quickly enough. This paper focuses on abort determination based primarily on data already available from the GN&C system. This work is the result of failure analysis efforts performed during the Ares I launch vehicle development program. Derivation of attitude and attitude rate abort triggers to ensure that abort occurs as quickly as possible when needed, but that false positives are avoided, forms a major portion of the paper. Some of the potential failure modes requiring use of these triggers are described, along with analysis used to determine the success rate of getting the crew off prior to vehicle demise.
NASA s next manned launch platform for missions to the moon and Mars are the Orion and Ares systems. Many critical aspects of the launch system performance are being verified using computational fluid dynamics (CFD) predictions. The Orion Launch Abort Vehicle (LAV) consists of a tower mounted tractor rocket tasked with carrying the Crew Module (CM) safely away from the launch vehicle in the event of a catastrophic failure during the vehicle s ascent. Some of the predictions involving the launch abort system flow fields produced conflicting results, which required further investigation through ground test experiments. Ground tests were performed to acquire data from a hot supersonic jet in cross-flow for the purpose of validating CFD turbulence modeling relevant to the Orion Launch Abort Vehicle (LAV). Both 2-component axial plane Particle Image Velocimetry (PIV) and 3-component cross-stream Stereo Particle Image Velocimetry (SPIV) measurements were obtained on a model of an Abort Motor (AM). Actual flight conditions could not be simulated on the ground, so the highest temperature and pressure conditions that could be safely used in the test facility (nozzle pressure ratio 28.5 and a nozzle temperature ratio of 3) were used for the validation tests. These conditions are significantly different from those of the flight vehicle, but were sufficiently high enough to begin addressing turbulence modeling issues that predicated the need for the validation tests.
As of Draft 4.0 of the CCT-REQ-1130 requirements document for CCP, ISS Crew Transportation and Services Requirements Document, specific language for the verification of the abort capability requirement, 3.3.1.4, was added. The abort capability requirement ensures that the CTS under dispersed conditions is always capable of aborting from a failed LV. The Integrated Aborts IPT was asked to author a memo for how this verification might be completed. The following memo dictates one way that this requirement and its verification could be met, but this is the not the only method.
A system, method, and apparatus for a vented launch vehicle adaptor (LVA) for a manned spacecraft with a "pusher" launch abort system are disclosed. The disclosed LVA provides a structural interface between a commercial crew vehicle (CCV) crew module/service module (CM/SM) spacecraft and an expendable launch vehicle. The LVA provides structural attachment of the module to the launch vehicle. It also provides a means to control the exhaust plume from a pusher-type launch abort system that is integrated into the module. In case of an on-pad or ascent abort, which requires the module to jettison away from the launch vehicle, the launch abort system exhaust plume must be safely directed away from critical and dangerous portions of the launch vehicle in order to achieve a safe and successful jettison.
The mathematical models and computer program which are used to implement LASS are described. The computer program provides for a simulation of boost to orbit and abort capability from boost trajectories to a prescribed target. The abort target provides a decision point for engine shutdown from which the vehicle coasts to the vicinity of the selected abort recovery site. The simulation is a six degree of freedom simulation describing a rigid body. The vehicle is influenced by forces and moments from nondistributed aerodynamics. An adaptive autopilot is provided to control vehicle attitudes during powered and unpowered flight. A conventional autopilot is provided for study of vehicle during powered flight.
Results from tests in the NASA/MSFC Trisonic Wind Tunnel on 0.004-Scale Orbiter and External Tank Force Models in Close Proximity (RTLS Abort Separation Conditions) are presented. The primary test objectives were to obtain data concerning proximity effects on the aerodynamic forces and moments experienced by Vehicle 2A Configuration Shuttle Orbiter and External Tank during an abort separation (Return to Launch Site) at a Mach number of 5. Additionally, data on orbiter control effectiveness during such an abort was obtained. Proximity effects were investigated for relative angles of incidence from minus 5 deg to plus 10 deg of the orbiter FRL with respect to the external tank centerline over a range of vertical and longitudinal displacements from the mated position to 2.5 tank diameters below and 3 tank diameters aft of the mated position.