Results for “Mission design center” — Page 8

NASA NTRS2014

Analyzing the Impacts of Natural Environments on Launch and Landing Availability for NASA's Eploration Systems Development Programs

NASA is developing new capabilities for human and scientific exploration beyond Earth orbit. Natural environments information is an important asset for NASA's development of the next generation space transportation system as part of the Exploration Systems Development Program, which includes the Space Launch System (SLS) and MultiPurpose Crew Vehicle (MPCV) Programs. Natural terrestrial environment conditions - such as wind, lightning and sea states - can affect vehicle safety and performance during multiple mission phases ranging from prelaunch ground processing to landing and recovery operations, including all potential abort scenarios. Space vehicles are particularly sensitive to these environments during the launch/ascent and the entry/landing phases of mission operations. The Marshall Space Flight Center (MSFC) Natural Environments Branch provides engineering design support for NASA space vehicle projects and programs by providing design engineers and mission planners with natural environments definitions as well as performing custom analyses to help characterize the impacts the natural environment may have on vehicle performance. One such analysis involves assessing the impact of natural environments to operational availability. Climatological time series of operational surface weather observations are used to calculate probabilities of meeting or exceeding various sets of hypothetical vehicle-specific parametric constraint thresholds.

Altino, Karen M.↗

NASA NTRS1998

Field Guide for Designing Human Interaction with Intelligent Systems

The characteristics of this Field Guide approach address the problems of designing innovative software to support user tasks. The requirements for novel software are difficult to specify a priori, because there is not sufficient understanding of how the users' tasks should be supported, and there are not obvious pre-existing design solutions. When the design team is in unfamiliar territory, care must be taken to avoid rushing into detailed design, requirements specification, or implementation of the wrong product. The challenge is to get the right design and requirements in an efficient, cost-effective manner. This document's purpose is to describe the methods we are using to design human interactions with intelligent systems which support Space Shuttle flight controllers in the Mission Control Center at NASA/Johnson Space Center. Although these software systems usually have some intelligent features, the design challenges arise primarily from the innovation needed in the software design. While these methods are tailored to our specific context, they should be extensible, and helpful to designers of human interaction with other types of automated systems. We review the unique features of this context so that you can determine how to apply these methods to your project Throughout this Field Guide, goals of the design methods are discussed. This should help designers understand how a specific method might need to be adapted to the project at hand.

Malin, Jane T.↗

NASA NTRS2016

Development of a High Performance, Low Profile Translation Table with Wire Feedthrough for a Deep Space CubeSat

NEAScout, a 6U cubesat and secondary payload on NASA's EM-1, will use an 85 sq m solar sail to travel to a near-earth asteroid at about 1 Astronomical Unit (about 1.5 x 10(exp 8) km) for observation and reconnaissance1. A combination of reaction wheels, reaction control system, and a slow rotisserie roll about the solar sail's normal axis were expected to handle attitude control and adjust for imperfections in the deployed sail during the 2.5-year mission. As the design for NEAScout matured, one of the critical design parameters, the offset in the center of mass and center of pressure (CP/CM offset), proved to be sub-optimal. After significant mission and control analysis, the CP/CM offset was accommodated by the addition of a new subsystem to NEAScout. This system, called the Active Mass Translator (AMT), would reside near the geometric center of NEAScout and adjust the CM by moving one portion of the flight system relative to the other. The AMT was given limited design space - 17 mm of the vehicle's assembly height-and was required to generate +/-8 cm by +/-2 cm translation to sub-millimeter accuracy. Furthermore, the design must accommodate a large wire bundle of small gage, single strand wire and coax cables fed through the center of the mechanism. The bend radius, bend resistance, and the exposure to deep space environment complicates the AMT design and operation and necessitated a unique design to mitigate risks of wire bundle damage, binding, and cold-welding during operation. This paper will outline the design constraints for the AMT, discuss the methods and reasoning for design, and identify the lessons learned through the designing, breadboarding and testing for the low-profile translation stages with wire feedthrough capability.

Few, Alex↗

NASA NTRS2016

Development of a High-Performance, Low-Profile Translation Table with Wire Feedthrough for a Deep Space CubeSat

NEAScout, a 6U cubesat and secondary payload on NASA's EM-1, will use an 85 sq m solar sail to travel to a near-earth asteroid at about 1 Astronomical Unit (about 1.5 x 10(exp 8) km) for observation and reconnaissance1. A combination of reaction wheels, reaction control system, and a slow rotisserie roll about the solar sail's normal axis were expected to handle attitude control and adjust for imperfections in the deployed sail during the 2.5-year mission. As the design for NEAScout matured, one of the critical design parameters, the offset in the center of mass and center of pressure (CP/CM offset), proved to be sub-optimal. After significant mission and control analysis, the CP/CM offset was accommodated by the addition of a new subsystem to NEAScout. This system, called the Active Mass Translator (AMT), would reside near the geometric center of NEAScout and adjust the CM by moving one portion of the flight system relative to the other. The AMT was given limited design space - 17 mm of the vehicle's assembly height-and was required to generate +/-8 cm by +/-2 cm translation to sub-millimeter accuracy. Furthermore, the design must accommodate a large wire bundle of small gage, single strand wire and coax cables fed through the center of the mechanism. The bend radius, bend resistance, and the exposure to deep space environment complicates the AMT design and operation and necessitated a unique design to mitigate risks of wire bundle damage, binding, and cold-welding during operation. This paper will outline the design constraints for the AMT, discuss the methods and reasoning for design, and identify the lessons learned through the designing, breadboarding and testing for the low-profile translation stages with wire feedthrough capability.

Few, Alex↗

NASA NTRS2006

A Holistic Approach for Risk Management During Design

In this paper, an approach for the identification, assessment, mitigation and continuous management of risks during the process of designing a space mission is presented. This approach has been developed by observing the risk patterns that occur at the Project Design Center of the Jet Propulsion Laboratory (TeamX) which develops conceptual, concurrent design of Space Missions. TeamX develops an end-to-end conceptual design of a Space Mission in a matter of one or two weeks. As the risk chair in TeamX, the author has had the opportunity to observe the risk patterns that occur during design over the course of many design sessions. This paper introduces an abstraction and generalization of those patterns. Risk is defined as anything that can go wrong, along with its approximate likelihood and consequence. The indicators, and causes, and effects of these risks are cross cutting across the multiple levels of people and processes involved in the design, and the actual design product itself.

reliability analysis↗

NASA NTRS1999

New frontiers in design synthesis

The Intelligent Synthesis Environment (ISE), which is one of the major strategic technologies under development at NASA centers and the University of Virginia, is described. One of the major objectives of ISE is to significantly enhance the rapid creation of innovative affordable products and missions. ISE uses a synergistic combination of leading-edge technologies, including high performance computing, high capacity communications and networking, human-centered computing, knowledge-based engineering, computational intelligence, virtual product development, and product information management. The environment will link scientists, design teams, manufacturers, suppliers, and consultants who participate in the mission synthesis as well as in the creation and operation of the aerospace system. It will radically advance the process by which complex science missions are synthesized, and high-tech engineering Systems are designed, manufactured and operated. The five major components critical to ISE are human-centered computing, infrastructure for distributed collaboration, rapid synthesis and simulation tools, life cycle integration and validation, and cultural change in both the engineering and science creative process. The five components and their subelements are described. Related U.S. government programs are outlined and the future impact of ISE on engineering research and education is discussed.

User-Computer Interface↗

NASA NTRS2003

A Distributed Simulation Software System for Multi-Spacecraft Missions

The paper will provide an overview of the web-based distributed simulation software system developed for end-to-end, multi-spacecraft mission design, analysis, and test at the NASA Goddard Space Flight Center (GSFC). This software system was developed for an internal research and development (IR&D) activity at GSFC called the Distributed Space Systems (DSS) Distributed Synthesis Environment (DSE). The long-term goal of the DSS-DSE is to integrate existing GSFC stand-alone test beds, models, and simulation systems to create a "hands on", end-to-end simulation environment for mission design, trade studies and simulations. The short-term goal of the DSE was therefore to develop the system architecture, and then to prototype the core software simulation capability based on a distributed computing approach, with demonstrations of some key capabilities by the end of Fiscal Year 2002 (FY02). To achieve the DSS-DSE IR&D objective, the team adopted a reference model and mission upon which FY02 capabilities were developed. The software was prototyped according to the reference model, and demonstrations were conducted for the reference mission to validate interfaces, concepts, etc. The reference model, illustrated in Fig. 1, included both space and ground elements, with functional capabilities such as spacecraft dynamics and control, science data collection, space-to-space and space-to-ground communications, mission operations, science operations, and data processing, archival and distribution addressed.

Burns, Richard↗

NASA NTRS2003

An Internet Protocol-Based Software System for Real-Time, Closed-Loop, Multi-Spacecraft Mission Simulation Applications

The paper will provide an overview of the web-based distributed simulation software system developed for end-to-end, multi-spacecraft mission design, analysis, and test at the NASA Goddard Space Flight Center (GSFC). This software system was developed for an internal research and development (IR&D) activity at GSFC called the Distributed Space Systems (DSS) Distributed Synthesis Environment (DSE). The long-term goal of the DSS-DSE is to integrate existing GSFC stand-alone test beds, models, and simulation systems to create a "hands on", end-to-end simulation environment for mission design, trade studies and simulations. The short-term goal of the DSE was therefore to develop the system architecture, and then to prototype the core software simulation capability based on a distributed computing approach, with demonstrations of some key capabilities by the end of Fiscal Year 2002 (FY02). To achieve the DSS-DSE IR&D objective, the team adopted a reference model and mission upon which FY02 capabilities were developed. The software was prototyped according to the reference model, and demonstrations were conducted for the reference mission to validate interfaces, concepts, etc. The reference model, illustrated in Fig. 1, included both space and ground elements, with functional capabilities such as spacecraft dynamics and control, science data collection, space-to-space and space-to-ground communications, mission operations, science operations, and data processing, archival and distribution addressed.

Davis, George↗

NASA NTRS2023

Distributed Spacecraft Mission (DSM) Plume Design Reference Mission (DRM) Inter-Satellite Link Modeling, Analysis and Simulation

NASA Goddard Space Flight Center (GSFC) Radical Innovation Initiative (RI2) plans to focus intently on DSM capability advancements in FY22-24. A DSM mission involves multiple spacecraft, arranged in a constellation, to achieve one or more common goals via the use of inter-satellite links (ISL) between the satellites. Recently, the GSFC Internal Research & Development (IRAD) program established Enceladus as a design reference mission (DRM) for the current DSM effort to foster the conceptual development of communication architecture, requirements, and solutions for future DSM ISL, as well as being able to push other research areas of interest. Enceladus is an icy moon of the planet Saturn. The DRM Enceladus mission concept involves a constellation of 24 small satellites, orbiting Enceladus around 100 km altitude in 3 planes, as observing nodes for science measurement. The mission science data will be sent back to Earth through a relay orbiting Saturn, using the constellation’s inter-satellite links. A QualNet/STK simulation model of the Relay and constellation ISL optical and RF links is developed for the design and optimization of the link and orbital parameters, as well as the inter-networking protocols. Delay Tolerant Networking (DTN) is utilized in the application layer modeling. This paper describes the plume DRM mission concept of an Enceladus constellation to relay science data to Earth and includes the proposed communication architecture and operation concepts. We present details of the QualNet/STK engineering model for this communication scenario to simulate the end-to-end data traffic through multiple layers (physical, data link, networking, transport and application). A link analysis for the constellation’s ISL, constellation to Relay and Direct to Earth (DTE) optical link is provided and discussed. The results of end-to-end traffic simulation for the data throughout/latency evaluation and assessment of the communication architecture are presented. The investigation of the concept of optical multiple access (OMA) for the Plume DRM is discussed. The modeling and simulation methodology developed in this paper is applicable to other DSMs in near Earth and deep space such as Earth-Moon L1/L2 and Lunar regions.

Distributed Spacecraft Mission (DSM) Inter-satelli↗

NASA NTRS2010

Orion Entry Performance-Based Center-of-Gravity Box

The Orion capsule is designed both for Low Earth Orbit missions to the ISS and for missions to the moon. 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 physics of atmospheric re-entry determine the capability of the Orion vehicle. For a given vehicle mass and shape, physics tells us that the driving parameters for an entry vehicle are the hypersonic lift-to-drag ratio (L/D) and the flight path angle at entry interface (gamma(sub EI)). The design of the Orion atmospheric re-entry must meet constraints during both nominal and dispersed flight conditions on landing accuracy, heating rate, total heat load, sensed acceleration, and proper disposal of the Service Module. These constraints define an entry corridor in the space of L/D-gamma(sub EI); if the vehicle falls within this corridor, then all constraints are met. The gamma(sub EI) dimension of the corridor can be further constrained by the gloads experienced during emergency entries. Thus, the entry performance for the Orion vehicle can be described completely by the L/D. Bounds on the hypersonic L/D necessary to achieve all the mission requirements can be defined for the given entry corridor. Landing accuracy performance drives the lower limit on L/D. In order to achieve the desired landing accuracy, a minimum L/D must be ensured. The design of the Thermal Protection System (TPS) drives the upper limit on L/D. A higher L/D can drive mass into the design of the TPS. Conversely, once the TPS is designed, the L/D must be ensured to stay below a certain limit in order for the TPS to stay within its design envelop. The L/D must stay within its upper and lower bounds during dispersed flight conditions. L/D is a function of both the aerodynamics and the center-of-gravity (CG) of the vehicle. The aerodynamics of the vehicle are determined by Computational Fluid Mechanics (CFD) and wind tunnel tests. However, the aerodynamics are not known precisely. Instead, an aerodynamic database has been developed where the aerodynamic coefficients are known to fall within a probabilistic band defined by upper and lower bounds. It is expected that the probabilistic band will shrink after the first missions are flown and real-world data is collected. Until that time, the Orion must be designed to the current aerodynamic database. Thus, for a given aerodynamic database with given uncertainties, the allowable range in L/D can be mapped to an allowable box for the CG location. The CG box is used to set requirements on the dispersions allowed for vehicle packaging and cargo storage. As the aerodynamic uncertainties decrease, the size of the CG box can increase. This paper discusses the technique used to map the minimum and maximum L/D bounds set by the entry performance requirements to the allowable dispersions in CG while accounting for aerodynamic uncertainties. The L/D is defined as the ratio of the lift force to the drag force. It is equivalent to the ratio of lift coefficient (C(sub L)) over drag coefficient (C(sub D)). C(sub L) and C(sub D) are functions of Mach number (M) and angle of attack (alpha). A Mach number of 25 is used as a measuring point of the hypersonic L/D. Variations in C(sub L), C(sub D) and alpha cause variations in L/D. Equation (1) shows the three contributions to the variation in L/D.

Rea, Jeremy R.↗

NASA NTRS2012

Optimizing Flight Control Software With an Application Platform

Flight controllers in NASA s mission control centers work day and night to ensure that missions succeed and crews are safe. The IT goals of NASA mission control centers are similar to those of most businesses: to evolve IT infrastructure from basic to dynamic. This paper describes Mission Control Technologies (MCT), an application platform that is powering mission control today and is designed to meet the needs of future NASA control centers. MCT is an extensible platform that provides GUI components and a runtime environment. The platform enables NASA s IT goals through its use of lightweight interfaces and configurable components, which promote standardization and incorporate useful solution patterns. The MCT architecture positions mission control centers to reach the goal of dynamic IT, leading to lower cost of ownership, and treating software as a strategic investment.

Smith, Irene Skupniewicz↗

NASA NTRS2020

Towards Crew-Centered, Mission-Oriented Space Flight Training

This poster describes a training approach that applies empirically derived principles of training to reimagining the overall design of NASA's space flight training program. The poster is focused specifically on the design of astronaut training for NASA's future deep space, exploration missions to Mars. We briefly describe NASA's space flight training practices during the Apollo and Space Shuttle eras as well as NASA's current practices for training astronauts for their missions to the International Space Station. We provide an overview of NASA's current concepts for a mission to Mars to scope our training approach. We envision a new space flight training approach which we term crew-centered, mission oriented training, inspired by the design approach offered in the context of airline pilot training by Barshi. We apply research-based training principles reviewed by Kole and his colleagues, as well as by other researchers in training science, into real-world, practical guidelines for the particular context of training astronauts for a mission to Mars.

Dempsey, Donna L.↗

NASA NTRS1992

Microwave systems applications in deep space telecommunications and navigation - Space Exploration Initiative architectures

The general support requirements of a typical SEI mission set, along with the mission operations objectives and related telecommunications, navigation, and information management (TNIM) support infrastructure options are described. Responsive system architectures and designs are proposed, including a Mars orbiting communications relay satellite system and a Mars-centered navigation capability for servicing all Mars missions. With the TNIM architecture as a basis, key elements of the microwave link design are proposed. The needed new technologies which enable these designs are identified, and current maturity is assessed.

Hall, Justin R.↗

NASA NTRS1984

Telecommunications and data acquisition support for the Pioneer Venus Project: Pioneers 12 and 13, prelaunch through March 1984

The support provided by the Telecommunications and Data Acquisition organization of the Jet Propulsion Laboratory (JPL) to the Pioneer Venus missions is described. The missions were the responsibility of the Ames Research Center (ARC). The Pioneer 13 mission and its spacecraft design presented one of the greatest challenges to the Deep Space Network (DSN) in the implementation and operation of new capabilities. The four probes that were to enter the atmosphere of Venus were turned on shortly before arrival at Venus, and the DSN had to acquire each of these probes in order to recover the telemetry being transmitted. Furthermore, a science experiment involving these probes descending through the atmosphere required a completed new data type to be generated at the ground stations. This new data type is known as the differential very long baseline interferometry. Discussions between ARC and JPL of the implementation requirements involved trade-offs in spacecraft design and led to a very successful return of science data. Specific implementation and operational techniques are discussed, not only for the prime mission, but also for the extended support to the Pioneer 12 spacecraft (in orbit around Venus) with its science instruments including that for radar observations of the planet.

Miller, R. B.↗

NASA NTRS1993

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.↗

NASA NTRS2000

Applying Multiagent Simulation to Planetary Surface Operations

This paper describes a multiagent modeling and simulation approach for designing cooperative systems. Issues addressed include the use of multiagent modeling and simulation for the design of human and robotic operations, as a theory for human/robot cooperation on planetary surface missions. We describe a design process for cooperative systems centered around the Brahms modeling and simulation environment being developed at NASA Ames.

Sierhuis, Maarten↗

NASA NTRS2013

SMAP Radar Processing and Calibration

The Soil Moisture Active Passive (SMAP) mission is part of the NASA space-based Earth observation program, and consists of an L-band radar and radiometer scheduled for launch into sun synchronous orbit in late 2014. A joint effort of the Jet Propulsion Laboratory (JPL) and the Goddard Space Flight Center (GSFC), the SMAP mission draws heavily on the design and risk reduction heritage of the Hydrosphere State (Hydros) mission [1], [2]. The SMAP science and applications objectives are to: 1) understand processes that link the terrestrial water, energy and carbon cycles, 2) estimate global water and energy fluxes at the land surface, 3) quantify net carbon flux in boreal landscapes, 4) enhance weather and climate forecast skill, and 5) develop improved flood prediction and drought monitoring capability. To meet these science objectives, SMAP ground processing will combine the attributes of the radar and radiometer observations (in terms of their spatial resolution and sensitivity to soil moisture, surface roughness, and vegetation) to estimate soil moisture with 4% volumetric accuracy at a resolution of 10 km, and freeze-thaw state at a resolution of 1-3 km. Model sensitivities translate the soil moisture accuracy to a radar backscatter accuracy of 1 dB (1 sigma) at 3 km resolution and a brightness temperature accuracy of 1.3 K at 40 km resolution. This paper will describe the level 1 radar processing and calibration challenges and the choices made so far for the algorithms and software implementation.

Synthetic Aperture Radar (SAR),↗

NASA NTRSDate not supplied

Quiet Spacecraft Cabin Ventilation Fan Development: Motivation and Context

It is important to control acoustical noise aboard crewed space vehicles and space habitats to provide a satisfactory environment for voice communications, alarm audibility, and restful sleep, and to minimize the risk for hearing loss and annoyance. As with most noise control efforts, it is best to control the noise at the source, and for spaceflight vehicles these are typically the fans associated with the Environmental Control and Life Support system. These include cabin ventilation fans and intermodule ventilation fans, among others. Throughout the history of crewed spaceflight, there have been issues with noise from ventilation fans, going back to Apollo. And, with the spaceflight vehicles and habitats currently being developed, there are again concerns with noise from ventilation fans. To address this problem, NASA is working to leverage the technology developed in its Aeronautics Research Mission Directorate, specifically at the Glenn Research Center, to design highly efficient and quiet fans for reducing community noise levels from civilian aircraft. This paper will discuss the need for transfer of this technology to spaceflight use, and will discuss the potential benefits, which are significant and will be crucial for the successful development and operations of future spaceflight vehicles, space suits, and habitats.

Christopher S. Allen↗