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NASA is working on the Constellation Program to go to the Moon and beyond. The Program contains a number of projects: a) The Ares I Launch Vehicle for the Orion Crew Module; b) The Ares V Launch Vehicle to launch large payloads into orbit, including the Earth Departure Stage (EDS); d) The Orion vehicle to transport the crew to orbit and to other destinations; and d - The Altair Lunar Lander
The following report was generated to give the International Space Station (ISS) Program some additional insight into the operations and issues associated with accommodating a single U.S. developed Assured Crew Return Vehicle (ACRV). During the generation of this report, changes in both the ISS and ACRV programs were factored into the analysis with the realization that most of the work performed will eventually need to be repeated once the two programs become more integrated. No significant issues associated with the ISS accommodating the ACRV were uncovered. Kinematic analysis of ACRV installation showed that there are viable methods of using Shuttle and Station robotic manipulators. Separation analysis demonstrated that the ACRV departure path clears the Station structure for all likely contingency scenarios. The payload bay packaging analysis identified trades that can be made between payload bay location, Shuttle Remote Manipulator System (SRMS) reach and eventual designs of de-orbit stages and docking adapters.
Under the NASA Exploration Technology Development Program (ETDP) and in partnership with the Department of Energy (DOE), NASA has embarked on a project to develop Fission Surface Power (FSP) technology. The primary goals of the project are to 1) develop FSP concepts that meet expected surface power requirements at reasonable cost with added benefits over other options, 2) establish a hardwarebased technical foundation for FSP design concepts and reduce overall development risk, 3) reduce the cost uncertainties for FSP and establish greater credibility for flight system cost estimates, and 4) generate the key products to allow NASA decision-makers to consider FSP as a preferred option for flight development. The FSP project was initiated in 2006 as the Prometheus Program and the Jupiter Icy Moons Orbiter (JIMO) mission were phased-out. As a first step, NASA Headquarters commissioned the Affordable Fission Surface Power System Study to evaluate the potential for an affordable FSP development approach. With a cost-effective FSP strategy identified, the FSP team evaluated design options and selected a Preliminary Reference Concept to guide technology development. Since then, the FSP Preliminary Reference Concept has served as a point-of-departure for several NASA mission architecture studies examining the use of nuclear power and has provided the foundation for a series of "Pathfinder" hardware tests. The long-term technology goal is a Technology Demonstration Unit (TDU) integrated system test using full-scale components and a non-nuclear reactor simulator. The FSP team consists of Glenn Research Center (GRC), Marshall Space Flight Center (MSFC) and the DOE National Laboratories at Los Alamos (LANL), Idaho (INL), Oak Ridge (ORNL), and Sandia (SNL). The project is organized into two main elements: Concept Definition and Risk Reduction. Under Concept Definition, the team performs trade studies, develops analytical tools, and formulates system concepts. Under Risk Reduction the team develops hardware prototypes and conducts laboratory-based testing.
Objectives for systems analysis capability: Develop integrated understanding of how a complex human physiological-socio-technical mission system behaves in spaceflight. Why? Support development of integrated solutions that prevent unwanted outcomes (Implementable approaches to minimize mission resources(mass, power, crew time, etc.)); Support development of tools for autonomy (need for exploration) (Assess and maintain resilience -individuals, teams, integrated system). Output of this exercise: -Representation of interfaces based on Human System Risk Board (HSRB) Risk Summary information and simple status based on Human Research Roadmap; Consolidated HSRB information applied to support communication; Point-of-Departure for HRP Element planning; Ability to track and communicate status of collaborations. 4
Following a curved path increases the distance to be flown, and a type of radio navigation that forces the adoption of such a path is therefore less efficient than one that marks out a definite straight line between the point of departure and the intended destination, and holds the airplane to that line. To determine the loss of efficiency resulting from curvature of the path, calculations were made for two particular cases by the method of step-by-step integration. The calculations were based on the assumption that the pilot makes straightforward use of his radio for navigation and makes no allowance for drift. Results are given in tabular form for two airplanes flying 200 miles at 100 mph, one with a cross wind of 50 mph wind across course, and the other with a 20 mph wind across course. It is shown that the following of the curved path increases the time of flight and the air distance flown by 17 percent and 2.5 percent in the two cases.
The International Space Station (ISS) will receive an armada of visiting supply vehicles during its life in orbit. Over 500 tons of material will be destroyed in targeted re-entries of these vehicles. Because all such re-entries lie in the same orbital plane of the station, and because the visiting vehicles typically deorbit within a few hours of departure, the ISS will usually be within sight of the re-entry process, at a range of only 300-600 kilometers. This vantage point offers an unprecedented opportunity for systematically measuring hypersonic destructive processes. This paper examines the integrated operational constraints of the ISS, its supply vehicles, and candidate sensors which can be employed in the scientific observation of the re-entry process. It is asserted the ISS program has the potential to reduce the worldwide risks from future deorbiting spacecraft, through systematic experimental characterization of the factors which affect the rupture, debris survival, and footprint size of its visiting vehicle fleet.
Additive Manufacturing (AM) has significantly evolved over the last decade for use in the aerospace industry, particularly for liquid rocket engines. AM offers a considerable departure from traditional manufacturing to rapidly fabricate components with complex internal features. High performance liquid rocket engine combustion chambers that operate in a high heat flux environment are fabricated using a copper-alloy liner with a series of integral coolant channels. Copper-alloys provide the necessary conductivity and material strength for adequate design margins offering high performance without the need for film coolant. Copper-alloys present unique challenges to properly melt the powder in laser-based AM processes due to their high reflectivity and conductivity. Starting in 2014, NASA’s Marshall Space Flight Center (MSFC) and Glenn Research Center (GRC) have developed a process for AM of GRCop (Copper- Chrome-Niobium) alloys using Selective Laser Melting (SLM). GRCop, originally developed at GRC, is a high conductivity, high-strength, dispersion strengthened copper-alloy for use in high-temperature, high heat flux applications. NASA has completed significant material characterization and testing, along with hot-fire testing, to demonstrate that GRCop-42 and GRCop-84 alloys are suitable for use in combustion chambers. Additional development and testing has been completed on AM bimetallic chambers using GRCop-84 liners and superalloy jackets, fabricated using two Directed Energy Deposition (DED) processes: Electron Beam Freeform Fabrication (EBF³) and Blown Powder DED. NASA completed hot-fire testing on various AM chambers using GRCop-84, GRCop-42, and bimetallic chambers in Liquid Oxygen (LOX)/Hydrogen, LOX/Methane, and LOX/Kerosene propellants.
Crewed missions to Mars present a challenging problem involving strong interactions between a variety of spacecraft systems. Design and architecture decisions made by one system can have powerful implications for another. Siloed subsystem development and optimization efforts are unlikely to generate optimal results for the integrated system and may lead to counterproductive investments at the mission and system level. Mismatched assumptions between different subsystems can also negatively impact the final result. Integrated systems analysis, trade studies, and sensitivity analysis are a critical element of successful Mars mission design. Previous work examined the linkages between Environmental Control and Life Support System (ECLSS) loop closure, food water content, spares Probability of Sufficiency (POS), and other factors by performing an integrated mass and cost assessment of a variety of Mars habitat and transportation system designs. This paper expands upon that work by evaluating the impact of mission duration, trajectory, and propulsion technology alongside key habitat variables. Transportation system and trajectory design decisions have enormous impacts on overall system mass, and the impacts of those decisions are themselves sensitive to the mass of the habitat that is being transported. This paper examines a variety of mission cases and evaluates total departure mass for each case. An estimate of relative costs for specific cases is also provided. The results are discussed, with an emphasis on the sensitivity of mass, cost, and other figures of merit to key parameters. Interactions and dependencies between various decisions, particularly those that relate to technology investments, are also highlighted. Overall, the various elements, systems, and subsystems required to support human spaceflight should not be optimized and evaluated in isolation, but instead should be analyzed as part of an integrated system. This paper demonstrates how integrated analysis can improve system understanding and lead to more optimal and effective crewed exploration systems.
Crewed missions to Mars present a challenging problem involving strong interactions between a variety of spacecraft systems. Design and architecture decisions made by one system can have powerful implications for another. Siloed subsystem development and optimization efforts are unlikely to generate optimal results for the integrated system and may lead to counterproductive investments at the mission and system level. Mismatched assumptions between different subsystems can also negatively impact the final result. Integrated systems analysis, trade studies, and sensitivity analysis are a critical element of successful Mars mission design. Previous work examined the linkages between Environmental Control and Life Support System (ECLSS) loop closure, food water content, spares Probability of Sufficiency (POS), and other factors by performing an integrated mass and cost assessment of a variety of Mars habitat and transportation system designs. This paper expands upon that work by evaluating the impact of mission duration, trajectory, and propulsion technology alongside key habitat variables. Transportation system and trajectory design decisions have enormous impacts on overall system mass, and the impacts of those decisions are themselves sensitive to the mass of the habitat that is being transported. This paper examines a variety of mission cases and evaluates total departure mass for each case. An estimate of relative costs for specific cases is also provided. The results are discussed, with an emphasis on the sensitivity of mass, cost, and other figures of merit to key parameters. Interactions and dependencies between various decisions, particularly those that relate to technology investments, are also highlighted. Overall, the various elements, systems, and subsystems required to support human spaceflight should not be optimized and evaluated in isolation, but instead should be analyzed as part of an integrated system. This paper demonstrates how integrated analysis can improve system understanding and lead to more optimal and effective crewed exploration systems.
The objectives of currently planned exploration efforts, as well as those further in the future, require significant advancements in propulsion technologies. The current Lunar exploration architecture has set goals and mission objectives that necessitate the use of new systems and the extension of existing technologies beyond present applications. In the near term, the majority of these technologies are the result of a need to apply high performing cryogenic propulsion systems to long duration in-space applications. Advancement of cryogenic propulsion to these applications is crucial to provide higher performing propulsion systems that reduce the vehicle masses; enhance the safety of vehicle systems and ground operations; and provide a path for In-situ Resource Utilization (ISRU).Use of a LOX/LH2 main propulsion system for Lunar Lander Descent is a top priority because more conventional storable propellants are far from meeting the performance needs of the current architecture. While LOX/LH2 pump feed engines have been used in flight applications for many years, these engines have limited throttle capabilities. Engines that are capable of much greater throttling while still meeting high performance goals are a necessity to achieving exploration goals. Applications of LOX/CH4 propulsion to Lander ascent propulsion systems and reaction control systems are also if interest because of desirable performance and operations improvements over conventional storable systems while being more suitable for use of in-situ produced propellants. Within the current lunar architecture, use of cryogenic propulsion for the Earth Departure Stage and Lunar Lander elements also necessitate the need for advanced Cryogenic Fluid Management technologies. These technologies include long duration propellant storage/distribution, low-gravity propellant management, cryogenic couplings and disconnects, light weight composite tanks and support structure, and subsystem integration. In addition to the propulsive and fluid management system technologies described, many component level technologies are also required to enable to the success if the integrated systems. The components include, but are not limited to, variable/throttling valves, variable position actuators, leak detectors, light weight cryogenic fluid pumps, sensor technology and others. NASA, partnering with the Aerospace Industry must endeavor to develop these, and other promising propulsion technologies, to enable the implements of the country's goals in exploration of the Moon, Mars and beyond.
NASA s long-range goal is focused upon human exploration of Mars. Missions to Mars will require campaigns of multiple launches to assemble Mars Transfer Vehicles in Earth orbit. Launch campaigns are subject to delays, launch vehicles can fail to place their payloads into the required orbit, and spacecraft may fail during the assembly process or while loitering prior to the Trans-Mars Injection (TMI) burn. Additionally, missions to Mars have constrained departure windows lasting approximately sixty days that repeat approximately every two years. Ensuring high reliability of launching and assembling all required elements in time to support the TMI window will be a key enabler to mission success. This paper describes an integrated methodology for analyzing and improving the reliability of the launch and assembly campaign phase. A discrete event simulation involves several pertinent risk factors including, but not limited to: manufacturing completion; transportation; ground processing; launch countdown; ascent; rendezvous and docking, assembly, and orbital operations leading up to TMI. The model accommodates varying numbers of launches, including the potential for spare launches. Having a spare launch capability provides significant improvement to mission success.
In order to develop a computer program able to model the frictional heating of metals in high pressure oxygen or nitrogen a number of additions have been made to the frictional heating model originally developed for tests in low pressure helium. These additions include: (1) a physical property package for the gases to account for departures from the ideal gas state; (2) two methods for spatial discretization (finite differences with quadratic interpolation or orthogonal collocation on finite elements) which substantially reduce the computer time required to solve the transient heat balance; (3) more efficient programs for the integration of the ordinary differential equations resulting from the discretization of the partial differential equations; and (4) two methods for determining the best-fit parameters via minimization of the mean square error (either a direct search multivariable simplex method or a modified Levenburg-Marquardt algorithm). The resulting computer program has been shown to be accurate, efficient and robust for determining the heat flux or friction coefficient vs. time at the interface of the stationary and rotating samples.
Incompressible three dimensional, turbulent boundary layer (3DTBL) experiments were simulated numerically by integrating the boundary layer equations together with an algebraic eddy viscosity turbulence model. For the flow treated, the downstream portion, where the crossflow was large, was not predicted with the present computational method; the flow was significantly influenced by elliptic flow field effects. Departures from the boundary layer concept are indicated. Calculations agreed reasonably well with the mean flow development up to separation. In one experiment the normal pressure gradients were found to be neligible in regions with large skewing and allowed testing turbulence models using the boundary layer equations. The simulation of this flow compared favorably with the experimental data throughout the flow field and suggested the applicability of algebraic eddy viscosity models for 3DTBLs.
The overall objective of this work is to study nucleate boiling heat transfer under microgravity conditions in such a way that while providing basic knowledge of the phenomena, it also leads to development of simulation models and correlations that can be used as design tools for a wide range of gravity levels. In the study a building block type of approach is used and both pool and low velocity flow boiling are investigated. Starting with experiments using a single bubble, the complexity of the experiments is increased to two or three inline bubbles, to five bubbles placed on a two-dimensional grid. Finally, experiments are conducted where a large number of prescribed cavities nucleate on the heater and when a commercial surface is used. So far experiments have been conducted at earth normal gravity and in the reduced gravity environment of the KC-135 aircraft whereas experiments on the space station are planned. Modeling/complete numerical simulation of the boiling process is an integral part of the total effort. Experiments conducted with single bubbles formed on a nucleation site microfabricated on a polished silicon wafer show that for gravity levels (g) varying from 1.5g(sub e) to 0.01g(sub e), the bubble diameter at departure varies approximately as (g(sub e)/g)(exp 1/2) and the growth period as (g(sub e)/g). When bubbles merge either inline or in a plane, the bubble diameter at departure is found to be smaller than that obtained for a single bubble and shows a weaker dependence on the level of gravity. The possible reason is that as the bubbles merge they create fluid circulation around the bubbles, which in turn induces a lift force that is responsible for the earlier departure of the bubbles. The verification of this proposition is being sought through numerical simulations. There is a merger of two inline, three inline, and several bubbles in a plane in the low gravity environment of the KC-135 aircraft. After merger and before departure, a mushroom type of bubble with several stems attached to the heater surface is clearly evident. Local heat fluxes during growth and departure of a single bubble were also measured. It was found that during most of the growth period of the bubble, generally the wall heat flux decreased with time because of the increased dry area under the bubble. However, the heat flux increased rapidly just prior to departure of the bubble because of the transient conduction into the cold liquid rushing to fill the space vacated by the bubble as the bubble base shrinks. The measured heat fluxes at various radial locations are found to be in qualitative agreement with the numerical predictions. Single bubble studies at earth normal gravity have also been performed on surfaces oriented at different angles to the gravitational acceleration with flow parallel to the surface. It is found that in all cases the bubbles slide along the surface before lift-off from the surface. The lift force generated as a result of the relative motion between the sliding bubbles and the imposed flow is found to play an important role when the normal force due to buoyancy is reduced. An experimental apparatus for the study of the bubble behavior with imposed flow under reduced gravity conditions has been developed and will soon be employed for experiments in the KC-135 aircraft.
The ULTOR(Registered TradeMark) Passive Pose and Position Engine (P3E) technology, developed by Advanced Optical Systems, Inc (AOS), uses real-time image correlation to provide relative position and pose data for spacecraft guidance, navigation, and control. Potential data sources include a wide variety of sensors, including visible and infrared cameras. ULTOR(Registered TradeMark) P3E has been demonstrated on a number of host processing platforms. NASA is integrating ULTOR(Registerd TradeMark) P3E into its Relative Navigation System (RNS), which is being developed for the upcoming Hubble Space Telescope (HST) Servicing Mission 4 (SM4). During SM4 ULTOR(Registered TradeMark) P3E will perform realtime pose and position measurements during both the approach and departure phases of the mission. This paper describes the RNS implementation of ULTOR(Registered TradeMark) P3E, and presents results from NASA's hardware-in-the-loop simulation testing against the HST mockup.
Since many of the larger launch vehicles are operated near their design limits during the ascent phase of flight to optimize payload to orbit, it often becomes necessary to verify that the vehicle will remain within certification limits during the ascent phase as part of the go/no-go review made prior to launch. This paper describes the approach used to predict Ares I-X launch vehicle structural air loads and controllability prior to launch which represents a distinct departure from the methodology of the Space Shuttle and Evolved Expendable Launch Vehicle (EELV) programs. Protection for uncertainty of key environment and trajectory parameters is added to the nominal assessment of launch capability to ensure that critical launch trajectory variables would be within the integrated vehicle certification envelopes. This process was applied by the launch team as a key element of the launch day go/no-go recommendation. Pre-launch assessments of vehicle launch capability for NASA's Space Shuttle and the EELV heavy lift versions require the use of a high-resolution wind profile measurements, which have relatively small sample size compared with low-resolution profile databases (which include low-resolution balloons and radar wind profilers). The approach described in this paper has the potential to allow the pre-launch assessment team to use larger samples of wind measurements from low-resolution wind profile databases that will improve the accuracy of pre-launch assessments of launch availability with no degradation of mission assurance or launch safety.
The Small Aircraft Transportation System Higher Volume Operations (SATS HVO) concept holds the promise for increased efficiency and throughput at many of the nations under-used airports. This concept allows for concurrent operations at uncontrolled airports that under today s procedures are restricted to one arrival or one departure operation at a time, when current-day IFR separation standards are applied. To allow for concurrent operations, SATS HVO proposes several fundamental changes to today's system. These changes include: creation of dedicated airspace, development of new procedures and communications (phraseologies), and assignment of roles and responsibilities for pilots and controllers, among others. These changes would affect operations on the airborne side (pilot) as well as the groundside (controller and air traffic flow process). The focus of this paper is to discuss some of the issues and potential problems that have been considered in the development of the SATS HVO concept, in particular from the ground side perspective. Reasonable solutions to the issues raised here have been proposed by the SATS HVO team, and are discussed in this paper.
NASA is planning a series of short and long duration human and robotic missions to explore the Moon and then Mars. The series of missions will begin with a new crew exploration vehicle (called Orion) that will initially provide crew exchange and cargo supply support to the International Space Station (ISS) and then become a human conveyance for travel to the Moon. The Orion vehicle will be mounted atop the Ares I launch vehicle for a series of pre-launch tests and then launched and inserted into low Earth orbit (LEO) for crew exchange missions to the ISS. The Orion and Ares I comprise the initial vehicles in the Constellation system of systems that later includes Ares V, Earth departure stage, lunar lander, and other lunar surface systems for the lunar exploration missions. These key systems will enable the lunar surface exploration missions to be initiated in 2018. The complexity of the Constellation system of systems and missions will require a communication and navigation infrastructure to provide low and high rate forward and return communication services, tracking services, and ground network services. The infrastructure must provide robust, reliable, safe, sustainable, and autonomous operations at minimum cost while maximizing the exploration capabilities and science return. The infrastructure will be based on a network of networks architecture that will integrate NASA legacy communication, modified elements, and navigation systems. New networks will be added to extend communication, navigation, and timing services for the Moon missions. Internet protocol (IP) and network management systems within the networks will enable interoperability throughout the Constellation system of systems. An integrated network architecture has developed based on the emerging Constellation requirements for Orion missions. The architecture, as presented in this paper, addresses the early Orion missions to the ISS with communication, navigation, and network services over five phases of a mission: pre-launch, launch from T0 to T+6.5 min, launch from T+6.5 min to 12 min, in LEO for rendezvous and docking with ISS, and return to Earth. The network of networks that supports the mission during each of these phases and the concepts of operations during those phases are developed as a high level operational concepts graphic called OV-1, an architecture diagram type described in the Department of Defense Architecture Framework (DoDAF). Additional operational views on organizational relationships (OV-4), operational activities (OV-5), and operational node connectivity (OV-2) are also discussed. The system interfaces view (SV-1) that provides the communication and navigation services to Orion is also included and described. The challenges of architecting integrated network architecture for the NASA Orion missions are highlighted.