Plume impingement during separation of a two stage space shuttle vehicle
Optimum separation trajectory analysis for two stage space shuttle configuration considering plume impingement effect
Fossler, I.↗
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A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions
The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.
Mercer, Carolyn R.↗
A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions
The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.
Mercer, Carolyn R.↗
Evaluation of some significant issues affecting trajectory and control management for air-breathing hypersonic vehicles
Horizontal takeoff airbreathing-propulsion launch vehicles require near-optimal guidance and control which takes into account performance sensitivities to atmospheric characteristics while satisfying physically-derived operational constraints. A generic trajectory/control analysis tool that deepens insight into these considerations has been applied to two versions of a winged-cone vehicle model. Information that is critical to the design and trajectory of these vehicles is derived, and several unusual characteristics of the airbreathing propulsion model are shown to have potentially substantial effects on vehicle dynamics.
Hattis, Philip D.↗
Application of CFE/POST2 for Simulation of Launch Vehicle Stage Separation
The constraint force equation (CFE) methodology provides a framework for modeling constraint forces and moments acting at joints that connect multiple vehicles. With implementation in Program to Optimize Simulated Trajectories II (POST 2), the CFE provides a capability to simulate end-to-end trajectories of launch vehicles, including stage separation. In this paper, the CFE/POST2 methodology is applied to the Shuttle-SRB separation problem as a test and validation case. The CFE/POST2 results are compared with STS-1 flight test data.
Pamadi, Bandu N.↗
Evaluating Lunar Descent and Landing Performance From a Near Rectilinear Halo Orbit Using Linear Covariance Resetting Techniques
Upcoming lunar programs are striving the achieve precision landing in a safe and robust manner. Various elements impact this mission objective ranging from on-orbit operations with ground station tracking to incorporating relative sensors with hazard detection and avoidance (HDA) to support the final approach and landing phase. Modeling the impacts of ground tracking, trajectory replanning, relative navigation sensors, and particularly a potential HDA system on the integrated closed-loop GN\&C system performance poses a unique challenge due to the complexity and interaction with multiple facets of the vehicle including the trajectory design, sensing hardware, navigation system, guidance and targeting, and the overall mission concept of operations. This paper outlines techniques to systematically analyze and compare the performance impacts of ground tracking and replanning and an HDA system where the onboard navigation errors are reset or uploaded from an external source and the vehicle's reference trajectory is regenerated requiring the system dispersions to also be reset to reflect this in-flight profile adjustment. To illustrate the application of these general techniques for analyzing the performance impacts due to incorporating these resetting events, they are demonstrated with a human lunar descent and landing scenario starting from a near rectilinear halo orbit (NRHO) until the vehicle precisely reaches its predetermined landing site on the lunar surface. Performance metrics such as inertial and relative navigation errors, trajectory dispersions, footprint dispersions, and propellant usage are provided.
GN&C↗
A open loop guidance architecture for navigationally robust on-orbit docking
The development of an open-hop guidance architecture is outlined for autonomous rendezvous and docking (AR&D) missions to determine whether the Global Positioning System (GPS) can be used in place of optical sensors for relative initial position determination of the chase vehicle. Feasible command trajectories for one, two, and three impulse AR&D maneuvers are determined using constrained trajectory optimization. Early AR&D command trajectory results suggest that docking accuracies are most sensitive to vertical position errors at the initial conduction of the chase vehicle. Thus, a feasible command trajectory is based on maximizing the size of the locus of initial vertical positions for which a fixed sequence of impulses will translate the chase vehicle into the target while satisfying docking accuracy requirements. Documented accuracies are used to determine whether relative GPS can achieve the vertical position error requirements of the impulsive command trajectories. Preliminary development of a thruster management system for the Cargo Transfer Vehicle (CTV) based on optimal throttle settings is presented to complete the guidance architecture. Results show that a guidance architecture based on a two impulse maneuvers generated the best performance in terms of initial position error and total velocity change for the chase vehicle.
Chern, Hung-Sheng↗
Optimal Control within the Context of Multidisciplinary Design, Analysis, and Optimization
Multidisciplinary design, analysis and optimization involves modeling the interactions of complex systems across a variety of disciplines. The optimization of such systems can be a computationally expensive exercise with multiple levels of nested nonlinear solvers running under an optimizer.The application of optimal control in project development often involves performing trajectory optimization for fixed vehicle designs or parametric sweeps across some key vehicle properties.This information is then relayed to the subsystem design teams who update their designs and relay some bulk characteristics back to the trajectory optimization procedure.This iteration is then repeated until the design closes.However, with increasing interest in more tightly coupled systems, such as electric and hybrid-electric aircraft propulsion and boundary layer ingestion, this process is prone to ignore subtle coupling between vehicle subsystem designs and vehicle operation on a given mission.Integrating trajectory optimization into a tightly coupled multidisciplinary design procedure can be computationally prohibitive, depending on the complexity of the subsystem analyses and the optimal control technique applied.To address these issues a new optimal control software tool, Dymos, has been developed.Dymos is built upon NASA's OpenMDAO software and can leverage its capabilities to efficiently compute gradients for the optimization and optimize complex models in parallel on distributed memory systems.This report provides some explanation into the numerical methods employed in Dymos and provides several use cases that demonstrate its performance on traditional optimal control problems and improvements ino techniques have been used extensively in recent decades to solve a variety of optimal control problems, typically in the form of aerospace vehicle trajectory optimization.
pseudospectral↗
An optimal terminal guidance method for continuous powered space vehicles.
Error correcting problem in space trajectory of continuous powered space vehicle, discussing optimal terminal guidance
OPTIMAL CONTROL↗
Terminal guidance for continuous powered space vehicles.
Error correcting problem in space trajectory of continuous powered space vehicle, discussing optimal terminal guidance
ERROR CORRECTING DEVICE↗
Trajectory design and navigation analysis for Cargo Transfer Vehicle proximity operations
The paper documents an investigation of alternative (or off-nominal) final approach trajectories and terminal guidance algorithms for Cargo Transfer Vehicle (CTV) automated rendezvous and proximity operation in the vicinity of Space Station Freedom (SSF). Ten candidate trajectory designs, representing a survey of all approaches that reasonably satisfy the mission constraints, are compared qualitatively and quantitatively. It is shown that several fundamentally different types of approach strategies are feasible in the CTV application and enjoy more favorable passive collision avoidance and plume impingement properties than the nominal. They are, however, generally more costly in propellant. The extent to which an alternative approach is more costly is determined largely by the abort (or specifically the 'hold' or 'wave-off') requirements. These approach techniques are verified in a software simulation using a parametrically defined candidate navigation system, and proximity sensor accuracy requirements are derived.
Marsh, S. M.↗
Real-Time Trajectory Assessment and Abort Management for the X-33 Vehicle
The X-33 is a flying testbed to evaluate technologies and designs for a reusable Single Stage To Orbit (SSTO) production vehicle. Although it is sub-orbital, it is trans-atmospheric. This paper will discuss the abort capabilities, both commanded and autonomous, available to the X-33. The cornerstone of the abort capabilities is the Performance Monitor (PM) and it's supporting software. PM is an on-board 3-DOF simulation, which evaluates the vehicle ability to execute the current trajectory. The Abort Manager evaluates the results from PM, and, when indicated, computes and implements an abort trajectory.
Moise, M. C.↗
Minimum propellant consumption round-trip trajectories to Mars for constant-thrust, constant-specific-impulse vehicles with optimum coasting periods
Minimum propellant consumption trajectories to Mars for constant-thrust, constant-specific impulse vehicles with optimum coasting periods
INTERPLANETARY TRAJECTORY↗
DUKSUP - A high thrust trajectory optimization code
Designing missions on expendable launch vehicles (ELV's) includes determining launch vehicle performance capabilities and trajectory characteristics over the range of mission requirements for suitable launch periods. This analysis depends on mathematically modeling both the launch vehicle and the mission requirements. Generally the result is a mathematical model that is described by an objective function to be optimized subject to an assortment of algebraic and dynamic constraints. About 30 years ago, engineers at Lewis Research Center (LeRC) undertook the task of creating a software code to solve 3D versions of such problems and based it on the Calculus of Variations/Optimal Control Theory. One of these codes, DUKSUP, has been in use at LeRC for nearly 25 years, during which time it has played an important role in a large number of studies and actual missions. Currently it is being used by about 12 analysts in the Center's Advanced Space Analysis Office (ASAO) to do mission design, feasibility studies, corroboration of contractor data and planning studies for the Space Exploration Initiative (SEI). Today, it is one of the few ELV mission analysis production codes based on variational methods. With future ELV missions in mind, ASAO is presently creating a new code to upgrade DUKSUP's capabilities.
Balkanyi, Leslie R.↗
Dynamic upper atmospheric force model on stabilized vehicles for a high-precision trajectory computer program
The upper atmosphere model draws heavily on the behavior of the earth's upper atmosphere which exhibits cyclic as well as irregular variations in density profile, temperature, pressure, and composition in unison with solar activities as deduced from the more recent land-based and satellite observations. The lift and drag model is designed specifically for inertially stabilized vehicles of the Mariner class, with possible extension to gravity gradient stabilized vehicles of the GEOS class. The model considers operation in the free molecular flow regimes with large Knudsen numbers. The vehicle is considered a composite structure with basic components having well-defined shapes, each with its own surface characteristics in terms of temperature, reflectivity, and accommodation of free stream molecules. The model takes into account both the calculation of precise aerodynamic force coefficients in terms of expansion of modified Bessel functions in speed ratios and angle of attack, and approximate force coefficients when the speed ratios approach infinity. Other considerations include specular and diffused reflectivity, shielding, and shadow effects.
Khatib, A. R.↗
Path Planning: Differential Dynamic Programming and Model Predictive Path Integral Control on VTOL Aircraft
This paper explores two optimal control approaches, widely used in robotics, to establish their viability as real-time trajectory planners for vehicle configurations envisioned for the emerging aviation sector of Urban Air Mobility (UAM). Differential Dynamic Programming (DDP) enables planning over highly nonlinear dynamics using second-order approximations along a nominal trajectory, and displays quadratic convergence to a local solution. Model Predictive Path Integral (MPPI) is a stochastic sampling-based algorithm that can optimize for general cost criteria, including potentially highly nonlinear formulations, and supports parallel computation through the use of modern GPU hardware. In this work, DDP and MPPI were implemented using model predictive control (MPC), and the results indicate they are able to successfully transition the aircraft over different flight envelopes and generate trajectories unique to UAM vehicles.
Differential Dynamic Programming↗
Preliminary Performance Analyses of the Constellation Program ARES 1 Crew Launch Vehicle
By the time NASA's Exploration Systems Architecture Study (ESAS) report had been released to the public in December 2005, engineers at NASA's Marshall Space Flight Center had already initiated the first of a series of detailed design analysis cycles (DACs) for the Constellation Program Crew Launch Vehicle (CLV), which has been given the name Ares I. As a major component of the Constellation Architecture, the CLV's initial role will be to deliver crew and cargo aboard the newly conceived Crew Exploration Vehicle (CEV) to a staging orbit for eventual rendezvous with the International Space Station (ISS). However, the long-term goal and design focus of the CLV will be to provide launch services for a crewed CEV in support of lunar exploration missions. Key to the success of the CLV design effort and an integral part of each DAC is a detailed performance analysis tailored to assess nominal and dispersed performance of the vehicle, to determine performance sensitivities, and to generate design-driving dispersed trajectories. Results of these analyses provide valuable design information to the program for the current design as well as provide feedback to engineers on how to adjust the current design in order to maintain program goals. This paper presents a condensed subset of the CLV performance analyses performed during the CLV DAC-1 cycle. Deterministic studies include development of the CLV DAC-1 reference trajectories, identification of vehicle stage impact footprints, an assessment of launch window impacts to payload performance, and the computation of select CLV payload partials. Dispersion studies include definition of input uncertainties, Monte Carlo analysis of trajectory performance parameters based on input dispersions, assessment of CLV flight performance reserve (FPR), assessment of orbital insertion accuracy, and an assessment of bending load indicators due to dispersions in vehicle angle of attack and side slip angle. A short discussion of the various customers for the dispersion results, along with results and ramifications of each study, are also provided.
Phillips, Mark↗
Interplanetary guidance systems requirements study. Volume 2 - Computer program descriptions. Part 5 - Nominal atmospheric entry trajectories, version 1
Computer program for simulating lifting vehicle atmospheric entry flight trajectory
Source record↗