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Tracking and data system support with multimission and high rate telemetry systems for the Mariner Mars 1969 project
The scientific objectives were considered for a Phobos/Deimos mission. The payloads for a minimum useful instrument complement were developed. The rationale for a sample return mission is discussed, along with the scientific constraints and requirements for the acquisition of samples.
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Space station study logic, commercial applications, total mission set space station system, mission requirements, free flyer missions, a baseline time phased mission set, and space systems architecture are discussed.
The Propulsion Integrated Vehicle Health Management (IVHM) Technology Experiment (PITEX) is a continuing NASA effort being conducted cooperatively by the NASA Glenn Research Center, the NASA Ames Research Center, and the NASA Kennedy Space Center. It was a key element of a Space Launch Initiative risk-reduction task performed by the Northrop Grumman Corporation in El Segundo, California. PITEX's main objectives are the continued maturation of diagnostic technologies that are relevant to second generation reusable launch vehicle (RLV) subsystems and the assessment of the real-time performance of the PITEX diagnostic solution. The PITEX effort has considerable legacy in the NASA IVHM Technology Experiment for X-vehicles (NITEX) that was selected to fly on the X-34 subscale RLV that was being developed by Orbital Sciences Corporation. NITEX, funded through the Future-X Program Office, was to advance the technology-readiness level of selected IVHM technologies within a flight environment and to begin the transition of these technologies from experimental status into RLV baseline designs. The experiment was to perform realtime fault detection and isolation and suggest potential recovery actions for the X-34 main propulsion system (MPS) during all mission phases by using a combination of system-level analysis and detailed diagnostic algorithms.
Guidance, Navigation, and Control (GN&C) is the measurement and control of spacecraft position, velocity, and attitude in support of mission objectives. This paper provides an overview of a preliminary design of the GN&C system of the Lunar Lander Altair. Key functions performed by the GN&C system in various mission phases will first be described. A set of placeholder GN&C sensors that is needed to support these functions is next described. To meet Crew safety requirements, there must be high degrees of redundancy in the selected sensor configuration. Two sets of thrusters, one on the Ascent Module (AM) and the other on the Descent Module (DM), will be used by the GN&C system. The DM thrusters will be used, among other purposes, to perform course correction burns during the Trans-lunar Coast. The AM thrusters will be used, among other purposes, to perform precise angular and translational controls of the ascent module in order to dock the ascent module with Orion. Navigation is the process of measurement and control of the spacecraft's "state" (both the position and velocity vectors of the spacecraft). Tracking data from the Earth-Based Ground System (tracking antennas) as well as data from onboard optical sensors will be used to estimate the vehicle state. A driving navigation requirement is to land Altair on the Moon with a landing accuracy that is better than 1 km (radial 95%). Preliminary performance of the Altair GN&C design, relative to this and other navigation requirements, will be given. Guidance is the onboard process that uses the estimated state vector, crew inputs, and pre-computed reference trajectories to guide both the rotational and the translational motions of the spacecraft during powered flight phases. Design objectives of reference trajectories for various mission phases vary. For example, the reference trajectory for the descent "approach" phase (the last 3-4 minutes before touchdown) will sacrifice fuel utilization efficiency in order to provide landing site visibility for both the crew and the terrain hazard detection sensor system. One output of Guidance is the steering angle commands sent to the 2 degree-of-freedom (dof) gimbal actuation system of the descent engine. The engine gimbal actuation system is controlled by a Thrust Vector Control algorithm that is designed taking into account the large quantities of sloshing liquids in tanks mounted on Altair. In this early design phase of Altair, the GN&C system is described only briefly in this paper and the emphasis is on the GN&C architecture (that is still evolving). Multiple companion papers will provide details that are related to navigation, optical navigation, guidance, fuel sloshing, rendezvous and docking, machine-pilot interactions, and others. The similarities and differences of GN&C designs for Lunar and Mars landers are briefly compared.
In the development of flight insulation systems for large cryogenic orbital storage (spray on foam and multilayer insulation), testing need include all environments that are experienced during flight. While large efforts have been expended on studying, bounding, and modeling the orbital performance of the insulation systems, little effort has been expended on the ground hold and ascent phases of a mission. Historical cryogenic in-space systems that have flown have been able to ignore these phases of flight due to the insulation system being within a vacuum jacket. In the development phase of the Nuclear Mars Vehicle and the Shuttle Nuclear Vehicle, several insulation systems were evaluated for the full mission cycle. Since that time there had been minimal work on these phases of flight until the Constellation program began investigating cryogenic service modules and long duration upper stages. With the inception of the Cryogenic Propellant Storage and Transfer Technology Demonstration Mission, a specific need was seen for the data and as such, several tests were added to the Cryogenic Boil-off Reduction System liquid hydrogen test matrix to provide more data on a insulation system. Testing was attempted with both gaseous nitrogen (GN2) and gaseous helium (GHe) backfills. The initial tests with nitrogen backfill were not successfully completed due to nitrogen liquefaction and solidification preventing the rapid pumpdown of the vacuum chamber. Subsequent helium backfill tests were successful and showed minimal degradation. The results are compared to the historical data.
This presentation is a summary of the Exploration EVA System Concept of Operations for the EVA Exploration Workshop.
The essential elements of the design process consist of the mission definition phase that provides the system requirements, the conceptual design, the preliminary design and finally the detailed design. Mission definition is performed largely by operations analysts in conjunction with the customer. The result of their study is handed off to the systems engineers for documentation as the systems requirements. The document that provides these requirements is the basis for the further design work of the design engineers at the Lockheed-Georgia Company. The design phase actually begins with conceptual design, which is generally conducted by a small group of engineers using multidisciplinary design programs. Because of the complexity of the design problem, the analyses are relatively simple and generally dependent on parametric analyses of the configuration. The result of this phase is a baseline configuration from which preliminary design may be initiated.
This white paper focuses on the scientific rationale and motivation for deploying phase change processes-based spacecraft thermal management systems for space missions [1,2]. Phase change processes need to be implemented in the design of various life support systems, sub-critical cryogenic and thermal control systems, taking advantage of the latent heat of phase change so as to significantly increase the heat transfer rate per unit fluid mass. Experimental data are critical to understanding gravitational effects on complex phase change processes using novel materials and working fluids. Experimental data can be acquired by either designing new testing platforms or performing experiments with simulant fluids on existing microgravity platforms. With the acquired data, theoretical flow physics models and numerical simulations can then be validated for use under long-duration microgravity conditions. The acquired data and validated models will enable the deployment of game-changing thermal management systems with remarkably improved efficiency and stability in heat acquisition, transport, and rejection for long-duration space missions.
The general purpose of this study is to aid in the evaluation and design of multi-sensor navigation schemes proposed for the orbiter. The scope of the effort is limited to the post-entry, energy management, and approach and landing mission phases. One candidate system based on conventional navigation aids is illustrated including two DME (Distance Measuring Equipment) stations and ILS (Instrument Landing System) glide slope and localizer antennas. Some key elements of the system not shown are the onboard IMUs (Inertial Measurement Units), altimeters, and a computer. The latter is programmed to mix together (filter) the IMU data and the externally-derived data. A completely automatic, all-weather landing capability is required. Since no air-breathing engines will be carried on orbital flights, there will be no chance to go around and try again following a missed approach.
Preflight activities and the Skylab mission support effort for the thruster attitude control system (TACS) are documented. The preflight activities include a description of problems and their solutions encountered in the development, qualification, and flight checkout test programs. Mission support effort is presented as it relates to system performance assessment, real-time problem solving, flight anomalies, and the daily system evaluation. Finally, the detailed flight evaluation is presented for each phase of the mission using system telemetry data. Data assert that the TACS met or exceeded design requirements and fulfilled its assigned mission objectives.
In this paper we describe pilots of a concurrent-engineering, formulation-phase design team demonstrated in spacecraft missions.
NASA’s next flag-ship mission - Europa Clipper, will embark on a journey to Jupiter’s icy moon Europa in 2024 to assess its environment and habitability with a highly capable spacecraft. Post Jupiter-Orbit-Insertion, the spacecraft will be commanded to perform intricate, yet meticulously planned Europa flybys to perform science investigations using a suite of instruments, while withstanding Jupiter’s harsh radiation environment. The success of this mission is dependent on a well-coordinated project and its elements such as the flight hardware and software, the ground support and mission operations teams, procedures and other cross-cutting elements. The Europa Clipper project needs to ensure that these elements are realized at a reasonable confidence level prior to launch and other mission critical events. System Validation test and analysis activities exercise and confirm the integrity of the system of all project elements in the expected flight environment with reasonable stressing conditions. These activities go beyond system design requirements verification and are driven by validation objectives that describe the end-to-end functional and operational capabilities required during nominal and off-nominal flight-like scenarios and critical events. The challenges associated with validating that the Europa Clipper project as a whole can function and perform correctly to meet the intended mission objectives with the as-delivered capabilities of all of its elements are daunting. This paper discusses the systematic methodology established in the early implementation phase of the Europa Clipper project for developing System Validation activities and their validation objectives, and addressing any validation-related challenges on the project. Approaches include decomposition of mission objectives using activity timelines in the Mission Design plan for developing nominal scenarios, use of fault trees for exploring off-nominal cases and system boundaries, and use of Model-based Systems Engineering (MBSE) tools for planning and prioritizing these activities.
Jupiter’s icy moon Europa is a prime target in our exploration of potentially habitable worlds beyond Earth. The combination of a subsurface liquid water layer in contact with a rocky seafloor may yield an ocean rich in the elements and energy needed for the emergence of life, and for potentially sustaining life through time. Europa may hold the clues to one of NASA’s long-standing quests – to determine whether or not we are alone in the universe. The Europa Clipper mission will characterize Europa’s habitability as the first step in the search for potential life at Europa by conducting approximately four dozen flybys. The project is preparing for a System Integration Review in November of 2021. This paper will summarize changes from the mission’s Preliminary Design Review (PDR) baseline, including science objectives, launch vehicle, interplanetary trajectory, science tour, flight system, payload and mission phases. Since PDR several of the Level 1 science requirements were descoped and the science requirement flow was realigned, in order to control cost growth. The realignment deemphasized science redundancy, which had resulted in an over-constrained science tour, in favor of a more targeted linkage to the primary responding instrument. Additionally, changes to the PDR flight system design baseline include changes to the sounding radar antenna design, the magnetometer architecture, and the heat redistribution system pumps. All changes have been incorporated into a reference science tour used to demonstrate that the requirements and design of the flight system and payload are mature enough to support flight builds.
Previous design efforts of a cost effective and reliable regenerative life support system (RLSS) provided the foundation for the characterization of organisms or 'biological processors' in engineering terms and a methodology was developed for their integration into an engineered ecological LSS in order to minimize the mass flow imbalances between consumers and producers. These techniques for the design and the evaluation of bioregenerative LSS have now been integrated into a lunar base reference mission, emphasizing the phased implementation of components of such a BLSS. In parallel, a designers handbook was compiled from knowledge and experience gained during past design projects to aid in the design and planning of future space missions requiring advanced RLSS technologies. The lunar base reference mission addresses in particular the phased implementation and integration of BLS parts and includes the resulting infrastructure burdens and needs such as mass, power, volume, and structural requirements of the LSS. Also, operational aspects such as manpower requirements and the possible need and application of 'robotics' were addressed.
This Project Plan for the Small Astronomy Satellite (SAS-D) is Goddard Space Flight Center's plan for accomplishing the project. It is issued in compliance with NASA Handbook NHB 7121.2 and NMI 7120.1, dated May 4, 1970, Subject: Approval and Implementation of Office of Space Science and Applications (OSSA) Research and Development (R&D) Projects. The Project Plan, as approved by GSFC executive management and NASA Headquarters, is the project manager's charter and defines his specific plan of action. System definition phase activities of the SAS-D Project will be a continuation of the analytical work that established the overall feasibility of the project. The system concept derived from that analytical work is defined in the Phase A (Mission Definition Phase) technical report. System definition phase activities will include detailed engineering analysis and preliminary breadboarding necessary to establish a single project approach from the feasible approaches. This document will be revised at the end of the system definition phase to provide greater detail on acquisition phase planning and to specify a specific system design.
Since our completing the NIAC Phase I NIAC on this Advanced Aerocapture System, NASA Langley Research Center has funded or supported a number of studies and code enhancements through its Center Innovation Fund (CIF) and NASA’s NSTGRO and Internship Programs to mature the analysis capabilities and quantify the merits of the MHD Aerocapture System technology. These efforts have resulted in a plug and play analysis capability for assessing MHD aerocapture system performance for arrival at many planetary bodies of interest. Our efforts have especially focused on the potential mass savings for improving the capacity for science observations at Neptune and Triton. A re-cent Forbes article published “‘Orbital mechanics is probably going to decide for us whether we go to Uranus or Neptune because we need to flyby Jupiter,’ said Kunio Sayanagi at Hampton University, Virginia, who also worked on the Neptune Odyssey proposal…. Exactly when a mission can be sent to Uranus, or Neptune, depends on the relative position of Jupiter, which can help give a spacecraft a gravitational slingshot. That drastically shortens the cruise phase.” [1] Since shortening the cruise phase is important for these science missions, any mass savings enabled by the MHD Aerocapture System could be reallocated to increasing Thermal Protection System mass to allow faster arrival speeds and/or for onboarding additional payloads such as science instruments, batteries, or propellant for conducting more science for longer durations in the desired orbits. The analysis steps and codes for conducting trades and sizing vehicles for aerocapture are as follows: Step 1 is to conduct aeroheating analysis using LAURA of the selected entry vehicle shape to identify locations on the forebody where ionization and flow velocity are sufficient for producing Lo-rentz forces. LAURA is a multiblock structured grid finite-volume CFD solver developed at the NASA Langley Research Center. [2] LAURA has been used for aerothermal analysis support of the entry, de-scent and landing (EDL) phase of interplanetary missions over the last three decades [3-7]. Step 2 is to port the LAURA results into CFDWARP to calcu-late electrical and thermal conductivities of ionized flow for sizing MHD patch system and calculating Lorentz forces needed for controls analysis. CFDWARP is a CFD code that uses advanced nu-merical methods that enable the simulation of the full coupling between the aerodynamics, the magne-tohydrodynamics, and the non-neutral plasma sheaths. CFDWARP has the unique capability to simulate efficiently the non-neutral sheaths (near the electrodes) in coupled form with the quasi-neutral bulk MHD flow [8-11]. Step 3 is to link re-sults from LAURA and CFDWARP into POST2 for calculating entry trajectories and comparing MHD control results with other aerodynamic control strategies. The Program to Optimize Simulated Tra-jectories II (POST2) is a generalized point mass, discrete parameter targeting and optimization pro-gram. POST2 provides the capability to target and optimize point mass trajectories for multiple pow-ered or un-powered vehicles near an arbitrary rotat-ing, oblate planet [12]. Step 4: TPS sizing was per-formed using the Fully Implicit Ablation and Ther-mal-response code (FIAT) tool which computes the transient one-dimensional thermal response and surface thermochemistry of a multilayer stackup of thermal protection, bonding, and structural materi-als subject to aeroheating on one surface [13]. The sizing and margining methodology used was based on the approach documented by Mahzari and Milos [14] for the dual-layered heatshield for extreme entry environment technology (DL-HEEET) TPS concept. TPS analysis utilizes trajectory information from POST2. Using this step-wise plug and play MHD Aerocapture performance assessment process, our analysis targets a Neptune aerocapture trajectory that will place the spacecraft in an observation orbit for Triton. [15]. Magnetohydrodynamic (MHD) control of a 4.5-meter diameter MSL-style capsule resulted in TPS mass savings of nearly 2000 kg when using an MHD system mass of under 200 kg. The flight path for a vehicle using the MHD control strategy has a much lower heat rate and heat load compared to the conventional aerodynamic aerocapture strategies known as bank angle con-trolled (BAC) and direct force controlled (DFC). Both BAC and DFC have heat rates significantly greater than 1500 W/cm2 typically used as an upper limit for PICA. Thus, DL-HEEET TPS concept was required for the BAC and DFC control strategies. However, considering the more benign environ-ments for the MHD case, additional TPS concepts with improved mass efficiency were also assessed. PICA was considered for the MHD controlled strat-egy since the maximum heat rate was well within the limits (<1500 W/cm2) of PICA. TPS sizing re-sulted in a significant mass reduction. The PICA layer for this sizing case was about 7.8 cm. As a point of reference, the Mars 2020 mission, which used this same PICA concept, had a PICA thickness of 3.18 cm [16]. The trajectories used for the TPS sizing originat-ed from the POST2 simulations. The current, I, to an electromagnet configuration can be manipulated to allow for active control of the vehicle. Manipula-tion of the current, I, changes the magnetic field, B, which affects the Lorentz force and therefore the MHD drag force on the vehicle. Our analysis in-cluded both open-loop and close-loop control. Closed-loop control will enable improved overall performance when taking into account mission level uncertainties, such as interplanetary delivery errors and atmospheric modeling uncertainties. The open-loop and closed-loop MHD control cases do not dip as deep into the atmosphere as the aerodynamic cases. Three types of aerodynamic-only approach-es are investigated: bank angle modulation (BAM), director force control (DFC), and Drag Modulated. BAM and DFC make use of vehicle aerodynamic angles to steer the vehicle. Thus, changing the aer-odynamic forces acting on the vehicle for control, aerodynamic drag modulated case requires a vary-ing drag area to modulate the drag force. The MHD drag modulated case modulates MHD generated drag force that adds to the aerodynamic drag. This higher atmospheric activation of drag forces by the MHD patch results in significantly less heat flux on the vehicle. The MHD technology will enable shorter cruise times and deceleration of larger payloads for increasing the capacity for science at the Ice Giants or for returning astronauts to Earth from cislunar space or from Mars. The purpose of this presentation is to provide more details about this work and to highlight plans for further research and development including a flight demonstration.