Simulating Attitudes and Trajectories of Multiple Spacecraft
A computer program called "42" simulates the attitudes and trajectories of multiple spacecraft flying in formation anywhere in the Solar System.
Stoneking, Eric↗
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Overview of Additively Manufactured TPS Proposed Flight Test and Earth Re-Entry Capsule Design
A flight mechanics overview is presented of an Earth flight test designed to investigate a novel, 3D printed thermal protection system (TPS) that is currently in development at NASA as part of the Additive Manufacturing of Thermal Protection Systems project. The project is pioneering a method to print a thermal protection system onto an entry vehicle forebody one layer at a time. This method reduces labor and complexity as compared to traditional manufacturing methods while increasing mission-dependent customization of through-depth materials properties. The flight test has three objectives. First, subject the forebody stagnation point of a capsule equipped with additively manufactured TPS (AMTPS) material to peak heat fluxes in excess of 100 $W/cm^2$. Second, capture in-flight data to enable flight reconstruction and AMTPS material thermal response model improvement. Third, recover the capsule with data storage and forebody AMTPS intact to enable post-flight inspection and analysis of AMTPS performance. The flight test trajectory is designed to achieve a peak stagnation point, cold-wall, entry heat flux of 135~$W/cm^2$. Flight mechanics simulations are performed using the Program to Optimize Simulated Trajectories II (POST2) and Monte-Carlo analysis yields statistical percentiles on vehicle performance at key points along the trajectory. Based on the flight mechanics analysis presented in this paper, a prototype capsule was designed, partially fabricated, and underwent preliminary component stress testing in preparation for fabrication of the flight unit capsule. The capsule outer mold line is a modified version of the heritage Mars Microprobe geometry. The capsule has a 0.356~m diameter, a 30~kg mass, and a hypersonic ballistic coefficient of 300~$kg/m^2$. Sensor selection is guided by flight dynamics simulations with the goal of resolving the re-entry heating pulse. On-board instrumentation include forebody and aftbody pressure sensors and thermocouples, a 9-axis IMU, a GPS receiver, and an Iridium satellite modem, all of which collect and store data throughout flight via on-board avionics systems. A two-stage parachute system is designed to decelerate the capsule to touchdown velocities that will not result in significant fracture or deformation of the charred AMTPS material at ground impact.
Flight Mechanics↗
NASA Advanced Concepts Office, Earth-To-Orbit Team Design Process and Tools
The Earth-to-Orbit Team (ETO) of the Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC) is considered the pre-eminent go-to group for pre-phase A and phase A concept definition. Over the past several years the ETO team has evaluated thousands of launch vehicle concept variations for a significant number of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Augustine Report, Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). The ACO ETO Team is called upon to address many needs in NASA s design community; some of these are defining extremely large trade-spaces, evaluating advanced technology concepts which have not been addressed by a large majority of the aerospace community, and the rapid turn-around of highly time critical actions. It is the time critical actions, those often limited by schedule or little advanced warning, that have forced the five member ETO team to develop a design process robust enough to handle their current output level in order to meet their customer s needs. Based on the number of vehicle concepts evaluated over the past year this output level averages to four completed vehicle concepts per day. Each of these completed vehicle concepts includes a full mass breakdown of the vehicle to a tertiary level of subsystem components and a vehicle trajectory analysis to determine optimized payload delivery to specified orbital parameters, flight environments, and delta v capability. A structural analysis of the vehicle to determine flight loads based on the trajectory output, material properties, and geometry of the concept is also performed. Due to working in this fast-paced and sometimes rapidly changing environment, the ETO Team has developed a finely tuned process to maximize their delivery capabilities. The objective of this paper is to describe the interfaces between the three disciplines used in the design process: weights and sizing, trajectory, and structural analysis. The tools used to perform such analysis are INtegrated Rocket Sizing (INTROS), Program to Optimize Simulated Trajectories (POST), and Launch Vehicle Analysis (LVA) respectively. The methods each discipline uses to streamline their particular part of the design process will also be discussed.
Waters, Eric D.↗
NASA Advanced Concepts Office, Earth-To-Orbit Team Design Process and Tools
The Earth-to-Orbit Team (ETO) of the Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC) is considered the pre-eminent "go-to" group for pre-phase A and phase A concept definition. Over the past several years the ETO team has evaluated thousands of launch vehicle concept variations for a significant number of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Augustine Report, Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). The ACO ETO Team is called upon to address many needs in NASA's design community; some of these are defining extremely large trade-spaces, evaluating advanced technology concepts which have not been addressed by a large majority of the aerospace community, and the rapid turn-around of highly time critical actions. It is the time critical actions, those often limited by schedule or little advanced warning, that have forced the five member ETO team to develop a design process robust enough to handle their current output level in order to meet their customer's needs. Based on the number of vehicle concepts evaluated over the past year this output level averages to four completed vehicle concepts per day. Each of these completed vehicle concepts includes a full mass breakdown of the vehicle to a tertiary level of subsystem components and a vehicle trajectory analysis to determine optimized payload delivery to specified orbital parameters, flight environments, and delta v capability. A structural analysis of the vehicle to determine flight loads based on the trajectory output, material properties, and geometry of the concept is also performed. Due to working in this fast-paced and sometimes rapidly changing environment, the ETO Team has developed a finely tuned process to maximize their delivery capabilities. The objective of this paper is to describe the interfaces between the three disciplines used in the design process: weights and sizing, trajectory, and structural analysis. The tools used to perform such analysis are INtegrated Rocket Sizing (INTROS), Program to Optimize Simulated Trajectories (POST), and Launch Vehicle Analysis (LVA) respectively. The methods each discipline uses to streamline their particular part of the design process will also be discussed.
Waters, Eric D.↗
Development of a system for transition characterization
The present system for real-time boundary layer transition detection and characterization encompasses aerodynamic simulations, sensor characteristics, and a hybrid analog/digital system for signal analysis, verification, and compression. By using artificial, intermittently laminar/turbulent signals, in conjunction with trajectory simulation, it becomes possible to conduct algorithm development which is independent of, and in advance of, sensor buildup. Attention is given to the illustrative case of the Pegasus launch vehicle's trajectory.
Bertelrud, Arild↗
Polar wind ion dynamics in the magnetotail
The circulation of polar wind ions from the high-latitude ionosphere to the plasma sheet is investigated. Single-particle trajectory simulations for the geomagnetic tail show, in addition to the breaking of adiabaticity, the possible creation of new high-altitude mirror points. This trajectory feature results from an intense parallel deceleration imparted by the magnetic field rotation during fast ExB transport. This centrifugal deceleration yields a critical parallel velocity which depends on the magnitude of the convection electric field and below which ions remain trapped inside the neutral sheet.
Delcourt, D. C.↗
Mars 2020 Navigation Performance
On February 18, 2021, the Perseverance rover and Ingenuity helicopter demonstration landed at Jezero Crater. The Entry, Descent, and Landing (EDL) architecture, largely the same used to land Curiosity at Gale Crater on August 6, 2012, required high-fidelity flight dynamics simulation with two independent tools to verify performance. The process for creating the EDL simulation using the Dynamics Simulator for Entry, Descent and Surface landing (DSENDS) tool will be discussed, along with its use for for independent verification of the EDL statistical analysis results and reference trajectory simulation. Analysis and usage details both in development and cruise, along with post-landing assessment of the prediction performance of the simulation, will also be discussed.
Wagner, Sean↗
Simulation and analysis of a geopotential research mission
Computer simulations were performed for a Geopotential Research Mission (GRM) to enable the study of the gravitational sensitivity of the range rate measurements between the two satellites and to provide a set of simulated measurements to assist in the evaluation of techniques developed for the determination of the gravity field. The simulations were conducted with two satellites in near circular, frozen orbits at 160 km altitudes separated by 300 km. High precision numerical integration of the polar orbits were used with a gravitational field complete to degree and order 360. The set of simulated data for a mission duration of about 32 days was generated on a Cray X-MP computer. The results presented cover the most recent simulation, S8703, and includes a summary of the numerical integration of the simulated trajectories, a summary of the requirements to compute nominal reference trajectories to meet the initial orbit determination requirements for the recovery of the geopotential, an analysis of the nature of the one way integrated Doppler measurements associated with the simulation, and a discussion of the data set to be made available.
Schutz, B. E.↗
An Expert System-Driven Method for Parametric Trajectory Optimization During Conceptual Design
During the early phases of engineering design, the costs committed are high, costs incurred are low, and the design freedom is high. It is well documented that decisions made in these early design phases drive the entire design's life cycle cost. In a traditional paradigm, key design decisions are made when little is known about the design. As the design matures, design changes become more difficult in both cost and schedule to enact. The current capability-based paradigm, which has emerged because of the constrained economic environment, calls for the infusion of knowledge usually acquired during later design phases into earlier design phases, i.e. bringing knowledge acquired during preliminary and detailed design into pre-conceptual and conceptual design. An area of critical importance to launch vehicle design is the optimization of its ascent trajectory, as the optimal trajectory will be able to take full advantage of the launch vehicle's capability to deliver a maximum amount of payload into orbit. Hence, the optimal ascent trajectory plays an important role in the vehicle's affordability posture yet little of the information required to successfully optimize a trajectory is known early in the design phase. Thus, the current paradigm of optimizing ascent trajectories involves generating point solutions for every change in a vehicle's design parameters. This is often a very tedious, manual, and time-consuming task for the analysts. Moreover, the trajectory design space is highly non-linear and multi-modal due to the interaction of various constraints. When these obstacles are coupled with the Program to Optimize Simulated Trajectories (POST), an industry standard program to optimize ascent trajectories that is difficult to use, expert trajectory analysts are required to effectively optimize a vehicle's ascent trajectory. Over the course of this paper, the authors discuss a methodology developed at NASA Marshall's Advanced Concepts Office to address these issues. The methodology is two-fold: first, capture the heuristics developed by human analysts over their many years of experience; and secondly, leverage the power of modern computing to evaluate multiple trajectories simultaneously and therefore enable the exploration of the trajectory's design space early during the pre- conceptual and conceptual phases of design. This methodology is coupled with design of experiments in order to train surrogate models, which enables trajectory design space visualization and parametric optimal ascent trajectory information to be available when early design decisions are being made.
Dees, Patrick D.↗
Enabling Parametric Optimal Ascent Trajectory Modeling During Early Phases of Design
During the early phases of engineering design, the costs committed are high, costs incurred are low, and the design freedom is high. It is well documented that decisions made in these early design phases drive the entire design's life cycle. In a traditional paradigm, key design decisions are made when little is known about the design. As the design matures, design changes become more difficult -- in both cost and schedule -- to enact. Indeed, the current capability-based paradigm that has emerged because of the constrained economic environment calls for the infusion of knowledge acquired during later design phases into earlier design phases, i.e. bring knowledge acquired during preliminary and detailed design into pre-conceptual and conceptual design. An area of critical importance to launch vehicle design is the optimization of its ascent trajectory, as the optimal trajectory will be able to take full advantage of the launch vehicle's capability to deliver a maximum amount of payload into orbit. Hence, the optimal ascent trajectory plays an important role in the vehicle's affordability posture as the need for more economically viable access to space solutions are needed in today's constrained economic environment. The problem of ascent trajectory optimization is not a new one. There are several programs that are widely used in industry that allows trajectory analysts to, based on detailed vehicle and insertion orbit parameters, determine the optimal ascent trajectory. Yet, little information is known about the launch vehicle early in the design phase - information that is required of many different disciplines in order to successfully optimize the ascent trajectory. Thus, the current paradigm of optimizing ascent trajectories involves generating point solutions for every change in a vehicle's design parameters. This is often a very tedious, manual, and time-consuming task for the analysts. Moreover, the trajectory design space is highly non-linear and multi-modal due to the interaction of various constraints. Additionally, when these obstacles are coupled with The Program to Optimize Simulated Trajectories [1] (POST), an industry standard program to optimize ascent trajectories that is difficult to use, it requires expert trajectory analysts to effectively optimize a vehicle's ascent trajectory. As it has been pointed out, the paradigm of trajectory optimization is still a very manual one because using modern computational resources on POST is still a challenging problem. The nuances and difficulties involved in correctly utilizing, and therefore automating, the program presents a large problem. In order to address these issues, the authors will discuss a methodology that has been developed. The methodology is two-fold: first, a set of heuristics will be introduced and discussed that were captured while working with expert analysts to replicate the current state-of-the-art; secondly, leveraging the power of modern computing to evaluate multiple trajectories simultaneously, and therefore, enable the exploration of the trajectory's design space early during the pre-conceptual and conceptual phases of design. When this methodology is coupled with design of experiments in order to train surrogate models, the authors were able to visualize the trajectory design space, enabling parametric optimal ascent trajectory information to be introduced with other pre-conceptual and conceptual design tools. The potential impact of this methodology's success would be a fully automated POST evaluation suite for the purpose of conceptual and preliminary design trade studies. This will enable engineers to characterize the ascent trajectory's sensitivity to design changes in an arbitrary number of dimensions and for finding settings for trajectory specific variables, which result in optimal performance for a "dialed-in" launch vehicle design. The effort described in this paper was developed for the Advanced Concepts Office [2] at NASA Marshall Space Flight Center
Holt, James B.↗
CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD) Ballistic Range Tests
A series of ballistic range tests were performed on a scaled model of the Supersonic Flight Demonstration Test (SFDT) intended to test the Supersonic Inflatable Aerodynamic Decelerator (SIAD) geometry. The purpose of these experiments were to provide aerodynamic coefficients of the vehicle to aid in mission and vehicle design. The experimental data spans the moderate Mach number range, $3.8-2.0$, with a total angle of attack ($alpha_T$) range, $10o-20o$. These conditions are intended to span the Mach-$alpha$ space for the majority of the SFDT experiment. In an effort to validate the predictive capabilities of Computational Fluid Dynamics (CFD) for free-flight aerodynamic behavior, numerical simulations of the ballistic range experiment are performed using the unstructured finite volume Navier-Stokes solver, US3D. Comparisons to raw vehicle attitude, and post-processed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data gives further insights into dynamic flow coupling leading to a potential mechanism for dynamic instability.
CFD modeling↗
Oxidation of Reinforced Carbon-Carbon Subjected to Hypervelocity Impact
Results of arc-jet tests conducted at the NASA Johnson Space Center (JSC) on Reinforced Carbon-Carbon (RCC) samples subjected to hypervelocity impact are presented. The RCC test specimens are representative of RCC component used on the Space Shuttle Orbiter. The objective of the arc jet testing was to establish the oxidation characteristics of RCC when hypervelocity projectiles, simulating meteoroid/orbital debris (MOD), impact the RCC material. In addition, analytical modeling of the increased material oxidation in the impacted area, using measured hole growth data, to develop correlations for use in trajectory simulations is also discussed.
Curry, Donald M.↗
Extensions of the space trajectories error analysis programs
A generalized covariance analysis technique which permits the study of the sensitivity of linear estimation algorithms to errors in a priori statistics has been developed and programed. Several sample cases are presented to illustrate the use of this technique. Modifications to the Simulated Trajectories Error Analysis Program (STEAP) to enable targeting a multiprobe mission of the Planetary Explorer type are discussed. The logic for the mini-probe targeting is presented. Finally, the initial phases of the conversion of the Viking mission Lander Trajectory Reconstruction (LTR) program for use on Venus missions is discussed. An integrator instability problem is discussed and a solution proposed.
Adams, G. L.↗
ASTP (SA-210) launch vehicle operational flight trajectory dispersion analysis, volume 1
In order to establish realistic deviation limits for the ASTP (SA-210) Launch Vehicle Operational Flight Trajectory, a dispersion analysis was conducted. The nominal trajectory prescribed for this analysis is the ASTP (SA-210) Launch Vehicle 500 Pound Launch Window Opening OT. The error sources considered are those associated with predictions of vehicle characteristics, vehicle systems performance, and flight environment. The nominal vehicle, the boost trajectory simulations, the error sources, the analytic procedures utilized, and the results are discussed. Launch vehicle guidance system inaccuracies were determined from the guidance error analysis. These data are composed of individual error source trajectory parameter dispersion envelopes.
Williams, N. D.↗
On the possibility of the determining the average mass composition near 10 to the 14th power eV through the solar magnetic field
The discovery of primary ultrahigh energy (UHE) gamma-rays has spawned plans for a new generation of air shower experiments with unprecedented directional resolution. Such accuracy permits observation of a cosmic ray shadow due to the solar disc. Particle trajectory simulations through models of the large scale solar magnetic field were performed. The shadow is apparent above 10 to the 15th power eV for all cosmic ray charges /Z/ 26; at lower energies, trajectories close to the Sun are bent sufficiently for this shadow to be lost. The onset of the shadow is rigidity dependent, and occurs at an energy per nucleus of approx. Z x 10 to the 13th power eV. The possibility of determining the average mass composition near 10 to the 14th power eV from 1 year's observation at a mountain altitude array is investigated.
Lloyd-Evans, J.↗
CPAS Preflight Drop Test Analysis Process
Throughout the Capsule Parachute Assembly System (CPAS) drop test program, the CPAS Analysis Team has developed a simulation and analysis process to support drop test planning and execution. This process includes multiple phases focused on developing test simulations and communicating results to all groups involved in the drop test. CPAS Engineering Development Unit (EDU) series drop test planning begins with the development of a basic operational concept for each test. Trajectory simulation tools include the Flight Analysis and Simulation Tool (FAST) for single bodies, and the Automatic Dynamic Analysis of Mechanical Systems (ADAMS) simulation for the mated vehicle. Results are communicated to the team at the Test Configuration Review (TCR) and Test Readiness Review (TRR), as well as at Analysis Integrated Product Team (IPT) meetings in earlier and intermediate phases of the pre-test planning. The ability to plan and communicate efficiently with rapidly changing objectives and tight schedule constraints is a necessity for safe and successful drop tests.
Englert, Megan E.↗
InSight Entry, Descent and Landing Pre-Flight Performance Predictions
On November 26, 2018, the Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) lander successfully touched down on the surface of Mars. Over its seven-plus year development, NASA Langley Research Center’s (LaRC) Program to Optimize Simulated Trajectories II (POST2) was used to assess the mission’s Entry, Descent and Landing (EDL) vehicle system performance against related requirements across the full range of possible environmental and spacecraft conditions. Much of the simulation code was derived from the Phoenix mission, for which this vehicle is very similar. The InSight six degree-of-freedom simulation included models for Mars atmosphere, gravity and digital elevation maps of the landing location. Additionally, vehicle specific aerodynamic, parachute, engine, navigation sensor, flight software and landing radar models were also included. A set of dispersions for each model, as well as for additional simulation input parameters, were also included in order to provide a statistical, Monte Carlo prediction of the EDL system performance. An overview of the pre-flight performance assessments completed, including the various simulation campaigns used, will be provided. Ultimately, this work was critical in the assessment of readiness for InSight launch. A brief description of the use of this simulation in support of flight operations is also discussed.
Robert W Maddock↗
Mars 2020 Perseverance Entry Descent and Landing Simulation Statistical Analysis and Operations
On February 18, 2021, the Perseverance rover and Ingenuity helicopter demonstra- tion landed at Jezero Crater. The Entry, Descent, and Landing (EDL) architecture, largely the same used to land Curiosity at Gale Crater on August 6, 2012, required high-fidelity flight dynamics simulation with two independent tools to verify per- formance. The process for creating the EDL simulation using the Dynamics Sim- ulator for Entry, Descent and Surface landing (DSENDS) tool will be discussed, along with its use for for independent verification of the EDL statistical analy- sis results and reference trajectory simulation. Analysis and usage details both in development and cruise, along with post-landing assessment of the prediction performance of the simulation, will also be discussed.
"O'Farrell, Clara"↗