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3 Technology Requirements for Future Space Missions The committeeâs charge requested that the benefits of using the Constellation System be compared with those of using âalternative implementation approaches,â meaning other technologies and launch vehicles. Some large payloads can only be accommodated by the Ares V rocket. However, even if the Ares V is used, some of the future space mission concepts evaluated by the committee (see Chapter 2 and Appendix B) would require advanced technologies that have not been fully developed. In some cases the technologies, such as advanced sensors, are required in order for a specific spacecraft to accomplish its mission and therefore would likely be developed as part of the flight program for that mission. However, the committee noted that several technologies were required for multiple mission concepts and were essentially âmission enabling,â meaning that the mission could not be accomplished without them. These include propulsion technologies that might allow an alternative to the use of a heavy-lift launch vehicle such as the Ares V and are applicable to multiple missions (for example, aerocapture, which can be used at Venus, Mars, Titan, and Neptune; and solar sails, which can be used for the Solar Polar Imager and Interstellar Probe missions). If NASA develops these technologies, an Ares V launch vehicle might not be required for these missions but might enhance them. However, there are also missions for which the advanced propulsion technologies are complementary to the Constellation System, such as use of the combination of conventional propulsion and aerocapture/aerobraking for orbit insertion and adjustment at the outer planets. This approach would maximize the mass fraction of the scientific payloads. Because the committee was asked to evaluate alternative implementation approaches, because advanced tech- nology development is so important to the majority of these missions, and because many of the technologies affect several mission concepts, the committee sought more information on the status of NASAâs overall technology development program for space science missions, particularly in the area of advanced in-space propulsion tech- nologies and deep space communications. Of the mission concepts that the committee considered, the Interstellar Probe, Neptune Orbiter with Probes, Solar Polar Imager, and Titan Explorer could all directly benefit from some form of in-space propulsion technology even if the Ares V is available. Although NASA had a vigorous in-space propulsion technology development program in the past, that pro- gram has been reduced in recent years. Many of the in-space propulsion development efforts that the committee reviewed were approaching flight-readiness testing but are now in the stage of being shut down due to lack of funding. The committee does not endorse any of these in-space technologies or development efforts. They are discussed in detail here in order to serve as an introduction to what in-space propulsion options may be required for many of these missions. However, it is clear from recent experience that NASA cannot afford all of them and 67
68 LAUNCHING SCIENCE BOX 3.1 Technology Readiness Levels NASA employs a technology readiness level (TRL) process to assess the flight readiness of a System Test, Launch given technology under development on the basis of & Operations TRL 9 its level of maturity (see Appendix D in this report). The TRL is rated on a scale of 1 to 9 depending on the degree of concept development, analytic and System/Subsystem TRL 8 empirical validation, and ultimately, flight validation Development (see Figure 3.1.1). TRL 7 TRL 9: ctual system âflight provenâ through suc- A cessful mission operations Technology Demonstration TRL 6 TRL 8: ctual system completed and âflight A qualifiedâ through test and demonstration (ground or space) TRL 5 Technology TRL 7: ystem prototype demonstration in a space S Development environment TRL 6: ystem/subsystem model or prototype dem- S TRL 4 onstration in a relevant environment (ground Research to Prove or space) Feasibility TRL 3 TRL 5: omponent and/or breadboard validation in C relevant environment TRL 4: omponent and/or breadboard validation in C Basic Technology TRL 2 Research laboratory environment TRL 3: nalytical and experimental critical formula A TRL 1 and/or characteristic proof-of-concept TRL 2: echnology concept and/or application for- T mulated FIGURE 3.1.1â Technology readiness level scale. TRL 1: asic principle observed and reported B SOURCE: Courtesy Figure 3.1.1.eps of NASA. SOURCE: NASA In-Space Propulsion Web site at http://www.grc.nasa.gov/WWW/InSpace/when.html. must choose carefully how and what to fund so as to expand its mission capabilities. The importance of a tightly focused technology program to support future advanced missions such as those that would be conducted using the Constellation System led the committee to recommend that NASA develop a comprehensive strategic plan for advanced propulsion system development (see the section below entitled âPropulsion System Technology Summaryâ). The overall technology requirements for the mission concepts studied in this report are summarized in Table 3.1. The technology readiness levels (TRLs) used by NASA to assess the flight readiness of a given technol- ogy are listed in Box 3.1. For a fuller description of these TRLs, see Appendix D.
TABLE 3.1â Correlation of Technology Needs with the Mission Concepts Analyzed by the Committee Technologies Enabling Alternatives or Enabling/Enhancing Technologies Enhancing Performance of Ares V Robotic Human Solar/ Next- Space Assembly Assembly Nuclear- Ares V Free-flying Tethered Generation Nuclear and and Solar Electric Upper Mission Concept Constellations Flight DSN Reactors Servicing Servicing Aerocapture Sails Propulsion Stage Advanced Compton Telescope (ACT) Advanced Technology Large-Aperture Space â â â Telescope (ATLAST) Dark Ages Lunar Interferometer (DALI) â â â 8-Meter Monolithic Space Telescope â â â Exploration of Near Earth Objects via the Crew â Exploration Vehicle Generation-X (Gen-X) â â Interstellar Probe â â â â â Kilometer-Baseline Far-Infrared/Submillimeter â âa â Interferometer Modern Universe Space Telescope (MUST) â â Neptune Orbiter with Probes â â ? â â Nuclear Neptune Orbiter with Probes â â â ? â â Palmer Quest âa â Single Aperture Far Infrared (SAFIR) Telescope ? â Solar Polar Imager â âb â â â Solar Probe 2 ? â â Stellar Imager â ? Super-EUSO (Extreme Universe Space Observatory) ? â â Titan Explorer âa â ? â â NOTE: Basic enabling technologies are those that exist in concept at present but whose development is needed for some of the missions. Technologies enabling alternatives to Ares are those whose availability would provide options for the missions currently conceived as being executed without requiring an Ares V. Human assembly and servicing would enhance the capabilities of these missions. Some of the missions, like the Advanced Compton Telescope, would not directly benefit from these technology developments. DSN, Deep Space Network; â?â signifies that missions may benefit, in the opinion of the committee. Aerocapture, solar sails, and solar/nuclear-electric propulsion all fall into the category of in-space propulsion technologies. aSignificant mission enhancement. bOnly if solar sail is implemented. 69
70 LAUNCHING SCIENCE Finally, the committee also learned that some of the missions anticipate upgrades to the Deep Space Network (DSN) that are not currently funded or planned. Many of the potential Constellation System mission concepts reviewed in this report would have a major impact on the DSN. At the very least, the development of gigapixel detector arrays for future telescopes would vastly increase the amount of data to be relayed to the ground. Because advanced in-space propulsion technology and deep space communications were identified as so important to so many of these mission concepts, they are discussed in greater detail in this chapter. IN-SPACE PROPULSION TECHNOLOGIES A number of advanced propulsion systems are being developed by NASAâs In-Space Propulsion Technology Program Office. Located at the NASA Glenn Research Center, this officeâs stated objective is to âdevelop in-space propulsion technologies that can enable or benefit near to mid-term NASA science missions by significantly reduc- ing travel times required for transit to distant bodies, increasing scientific payload capability or reducing mission costs.â The In-Space Propulsion Technology Office primarily focuses on technology development and demonstra- tion, which represents a TRL range from 3 to 6. This provides the foundation for flight system development and technology implementation by the various science missions, including several of the missions reviewed in this study. Current development efforts include electric propulsion, solar sails, aerocapture, and advanced chemical propulsion. A brief overview of electric propulsion, solar sails, and aerocapture and their development status is provided below. Electric Propulsion Electric propulsion technologies generate thrust by using electrical energy to accelerate an onboard propellant, such as xenon gas, to very high ejection velocities. These velocities may reach 20 times those of traditional chemical propulsion systems. This is characterized by the specific impulse (abbreviated as Isp), which is the product of the thrust and firing duration of the engine, divided by the mass of the propellant expended. This increased efficiency means that less propellant needs to be carried onboard for a given mission. Electrical propulsion systems are subdivided by the sources of the electrical energy, such as solar electric propulsion (SEP), nuclear electric propulsion (NEP), and radioisotope electric propulsion (REP). Solar Electric Propulsion Solar electric propulsion is a technology that uses electrical energy from solar arrays to accelerate a propel- lant to very high velocities and obtained thrust. In its simplest form, SEP consists of a propellant supply, power processing unit (PPU), and an ion thruster (see Figure 3.1). NASA has already flown SEP successfully, and the Dawn spacecraft currently heading toward a rendezvous with the asteroid Vesta employs this technology. The Titan Explorer, Solar Polar Imager, and Neptune Orbiter with Probes all have mission design options that include the use of SEP. NASAâs Evolutionary Xenon Thruster.â NASAâs Evolutionary Xenon Thruster (NEXT) project builds on the technology developed for the NASA SEP Technology Application Readiness (NSTAR) system that served as the primary ion propulsion system for the Deep Space 1 (DS-1) spacecraft launched in 1998. The NSTAR system used gridded ion thrusters. During the DS-1 mission, the NSTAR engine operated for a total of 16,246 hours, or 6,477 days, easily surpassing the previous record of 161 days established by NASAâs Space Electric Rocket Test 2 (SERT 2), which flew in 1970. â Much of the information in this chapter was provided in a briefing to the committee during its June 2008 meeting: Tibor Kremic, âNASA Investment in In-Space Propulsion Technologies,â presentation to the Committee on Science Opportunities Enabled by NASAâs Constellation System, June 9-11, 2008. â Rae Mayer, âIn-Space Propulsion Technology Program Office (ISPT) Overview,â February 2006.
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 71 FIGURE 3.1â Electrostatic ion thruster. SOURCE: Courtesy of NASA. As does NSTAR, the NEXT system uses a gridded ion thruster, which uses two oppositely charged grids to accelerate and eject high-speed ions. The NEXT project incorporates advances in the gridded ion thruster, PPU, and propellant management system designs for increased system performance and life. It is a 7-kilowatt (kW) system, compared with the 2.3-kW NSTAR system. The NEXT thrusters have higher efficiency and operate with a higher Isp over a broader throttling range compared with the NSTAR thrusters. The NEXT thrusters also have increased propellant throughput, which reduces the number of thrusters required for equivalent mission require- ments. A comparison of the state of the art and NEXT thrusters is provided in Table 3.2. The PPU and propellant management system are similarly designed to be highly efficient, with lower specific mass. The propellant man- agement system provides tighter propellant flow-control capability, reducing the propellant residuals at the end of the mission and increasing overall mission performance. The first phase of the NEXT project was completed in 2003, during which a breadboard PPU, a breadboard propellant management system, and engineering model thrusters were built, and their component and system- level testing were completed (see Figure 3.2). In 2005, three engineering-model thrusters were tested as an array, TABLE 3.2â Comparison of Current State-of-the-Art Thrusters and NASAâs Evolutionary Xenon Thruster (NEXT), Under Development Thruster Attribute State of the Art NEXT Maximum input power (kW) 2.3 Up to 6.9 Throttle range 4:1 >12:1 Isp (s) 3,170 4,190 Efficiency (%) 62 71 Propellant throughput (kg) 150 >300 Specific weight (kg/kWe) 3.6 1.8
72 LAUNCHING SCIENCE FIGURE 3.2â NASAâs Evolutionary Xenon Thruster (NEXT), a gridded ion thruster. SOURCE: Courtesy of NASA. and an extended-duration test of a single engineering-model thruster was initiated. The extended-duration test exceeded 15,000 hours of operation with a 325 kg throughput. (See Figure 3.3.) Knowledge gained from the tests was incorporated into the prototype-model (PM-1) thruster and engineering-model PPU and propellant manage- ment system for the current phase of development. With the exception of the life testing, the NEXT systems have completed all component and system-level testing, achieving TRL 6. The life testing is in process and will continue through 2010. High Voltage Hall Accelerator Thruster.â The High Voltage Hall Accelerator (HiVAC) Thruster project is intended to develop a low-power, long-life Hall thruster. In contrast to the gridded ion thrusters, where two oppositely charged grids are used to accelerate and eject ions to generate thrust, Hall thrusters use a radial magnetic field to accelerate low-density plasma to high velocities. The Isp of Hall thrusters is generally in the range of 1,000 to 3,000 seconds (s). They can operate over a wide range of input power, allowing them to be used far from the Sun, where little solar energy is available. The major advantage of Hall thrusters is the potential of lower costs compared with costs for the gridded ion thrusters. Recent advances include the design and fabrication of a flight-prototype thruster, which is intended to be able to operate for more than 7,500 hours, with a specific impulse of 2,800 s at a maximum power level of 3.5 kW. A laboratory version has been designed to operate up to 15,000 hours under similar conditions. Thus far, 4,100 s of operation with 88-kg throughput have been demonstrated. At the 3.5-kW design power level, 55 percent total efficiency has been demonstrated. The design and fabrication of a full-engineering-model Hall thruster is to be completed this year. At this time, NASA has not funded the development of other system elements such as the PPU or propellant management system. Energetics Program Development.â Until 2005, the Energetics Program was funded to develop SEP technologies to TRL 3, setting the stage for future development. Areas of interest included electrostatic ion thruster technology, â The NASA Web page is available at http://www.grc.nasa.gov/WWW/RT/2006/RP/RPP-manzella.html.
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 73 FIGURE 3.3â NASA-94M, a state-of-the-art ion thruster. SOURCE: Courtesy of NASA. electrostatic Hall-effect thruster technology, pulsed-plasma thruster technology, high-power electric propulsion technology, and electric propulsion systems. This effort primarily focused on longer life, higher reliability, and higher power and thrust. Nuclear Electric Propulsion Missions to the outer planets drive the need for high-power, high-specific-impulse propulsion systems. Such missions include the Neptune Orbiter with Probes and the Interstellar Probe. Nuclear electric propulsion has the greatest potential to meet the propulsion needs of these missions. In addition to the high power levels, nuclear power plants provide a constant source of electrical energy throughout the mission, which is in direct contrast to SEP systems, for which the available solar energy declines dramatically as a spacecraft travels to the outer planets. NASAâs Prometheus program was developing NEP systems as a building block to space exploration. The initial mission focus was the study of three ice-covered moons of Jupiter: Ganymede, Callisto, and Europa. It was intended to study the vast oceans of water under the icy surface of the moons, looking for indications of or opportunities for life. The Jupiter Icy Moons Orbiter (JIMO) spacecraft (see Figure 3.4) was to be propelled by high-power electric propulsion (HiPEP) ion propulsion thrusters, with the electrical power supplied by a nuclear fission reactor. The reactor was to be positioned at the tip of a boom extended away from the spacecraft, with a strong radiation shield protecting the electronics and sensors on the spacecraft. As described in Box 1.2 in Chapter 1 of this report, due to a shift in NASA priorities and significant technical challenges associated with the JIMO mission, as well as the high costs of the reactor program, funding was cut, ending the Prometheus program. In support of the Prometheus program and future NEP systems, the In-Space Propulsion Technology Program Office was developing high-energy thruster technologies. The challenge for NEP system design is the increased power, performance, and life requirements of the thrusters. The NSTAR and NEXT thrusters developed for SEP systems provide 2.3 kW and 6.9 kW, respectively, compared with power requirements in excess of 20 kW for deep space missions. Similarly, the Isp requirement of 7,000+ s is significantly greater than the 3,170 s and 4,190 s pro- vided by NSTAR and NEXT. And, finally, where the NSTAR and NEXT thrusters demonstrated more than 15,000 hours of operational life, deep space missions will require in excess of 70,000 hours. Within the past 5 years, development efforts were completed for the HiPEP thruster and Nuclear Electric Xenon Ion System (NEXIS). The goal of NEXIS was to demonstrate a 20-kW system and a 2,000-kg xenon throughput providing an Isp of 6,000 s to 8,000 s, with a 7- to 11-year life. The HiPEP thruster was intended to provide a 25-kW system with a 2,000-kg xenon throughput and similar performance characteristics. A comparison of the objectives of NEXIS and HiPEP thrusters with the NSTAR and NEXT thruster designs is provided in Table 3.3.
74 LAUNCHING SCIENCE FIGURE 3.4â Artistâs concept of the nuclear-powered Jupiter Icy Moons Orbiter. SOURCE: Courtesy of NASA. Both development efforts advanced the technology to the TRL 3 to 4 range, laying the foundation for future high-power, long-life thruster development, including the thrusters for the JIMO mission. The programs each demonstrated performance and preliminary life characteristics but require additional component, subsystem, and system verification and validation through modeling and testing to bring the technology to TRL 6. TABLE 3.3â Comparison of Several High-Power Electric Propulsion Systems Under Development at NASA Performance Metric NSTAR NEXT NEXIS HiPEP Power (kWe) 2.3 7 20 23.5 Isp (s) 3,170 4,190 7,500 6,000-8,000 Thruster efficiency (%) 63 71 78 >74 Specific mass (kg/kWe) 3.6 1.8 1.5 <2 Throughput (kg) 150 >300 2,000 5,100 Life capability (thousand hr) 30 >15 93 243 NOTE: NSTAR, NASA SEP Technology Application Readiness; NEXT, NASAâs Evolutionary Xenon Thruster; NEXIS, Nuclear Electric Xenon Ion System; HiPEP, high-power electric propulsion.
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 75 FIGURE 3.5â A solar sail design during ground test. SOURCE: Courtesy of NASA. Solar Sails The Interstellar Probe and Solar Polar Imager mission design options included the use of solar sails. These devices use photons from the Sun, which apply pressure on thin, lightweight, highly reflective sheets or sails to propel the spacecraft, much like wind pushes a sailboat across a body of water (see Figure 3.5). The direction of the spacecraft is controlled by changing the angle of the solar sail relative to the Sun and spacecraft. Since the Sun provides all of the propulsive energy, there is no need for onboard propellants. Although the thrust associated with solar sail propulsion is very smallâorders of magnitude less than traditional chemical propulsion systemsâthe continuous application of solar energy can increase the spacecraft velocity to levels in excess of chemical propul- sion systems. The Solar Polar Imager concept uses a solar sail to approach and then increase the orbital inclination about the Sun, while the Interstellar Probe uses the technology to accelerate the vehicle to the edge of the solar system. Solar sails are most effective within 2 to 3 AU of the Sun. Since solar pressure is very low, sails that are very large and lightweight are required to produce appreciable thrust. The Jet Propulsion Laboratory studied the use of solar sails in the 1970s and determined that a 600,000-m 2 sail would be required to rendezvous with Halleyâs Comet, matching both its position and velocity. These very large sails must be stowed during launch and safely deployed once the vehicle has left Earthâs atmosphere. The support structures, or trusses, for the sail must also be very lightweight and need to be stowed in a small volume for launch. The sail material requires a highly reflective coating on the Sun-facing side in order to maximize the thrust, while retaining a highly emissive coating on the back side. The sail materials developed to date are 40 to 100 times thinner than a piece of standard writing paper. It is generally a durable polymer material, such as Mylar or Kapton, with a metalized surface to obtain the desired reflectivity. â See http://www.nasa.gov/vision/universe/roboticexplorers/solar_sails.html.
76 LAUNCHING SCIENCE In 2004 and 2005, ATK Space Systems and LâGarde, Inc., designed, fabricated, and tested 10-m and 20-m solar sails under thermal vacuum conditions for NASAâs In-Space Propulsion Technology Program Office. ATK was able to demonstrate greater than 90 percent reflectance across the solar spectrum and a backside emissivity of 0.3 or better, achieving the goals of the development program. Given the results of their study, ATK projected the area density and stowed-volume requirements to be met for the 100-m sail configuration. In addition, the com- pany projected that emissivities of greater than 0.5 can be achieved with carbon-loaded clear polyimide 1 (CP1), a temperature-resistant material invented by the NASA Langley Research Center specifically for space applications. Similar work by LâGarde, Inc., demonstrated 86 percent reflectance, slightly below the 90 percent goal, and an emissivity of 0.4, with a potential for 0.62 with a thicker deposition of blackened chromium. Both projects dem- onstrated robust attitude control, with the thrust turning rate significantly greater than design goals. NASA terminated the development effort at TRL 5 because of funding limitations. Additional work remains in order to develop improved, automated manufacturing capabilities, characterize material capabilities, and model and test the materials and lightweight, deployable structures. Improved, high-fidelity ground simulators are also required to offset the effect of gravity on the large sails. In addition, in order to meet the requirements of the Interstellar Probe mission, an order-of-magnitude improvement in area density of the sails would be needed. NASA attempted to demonstrate the deployment and propulsion capability of a solar sail in orbit. The Nano- Sail-D was intended to be the first spacecraft to use a solar sail for attitude control. The sail is a four-segment square, approximately 10 ft on a side. Unfortunately, the launch vehicle, the SpaceX Falcon 1 booster, failed at the August 2008 launch during the boost phase, and the mission was lost. In June 2005, Cosmos Studios and the Planetary Society had also attempted to test a solar sail in space using a submarine-launched missile to place the Cosmos 1 spacecraft into Earth orbit. The Cosmos 1 concept employed eight triangular sails, each 15 m in length, to provide a total sail area of 600 m2. Unfortunately the launch vehicle failed, preventing Cosmos 1 from reach- ing orbit. The Planetary Society is working on Cosmos 2, which is intended to make a controlled flight using solar pressure. The society hopes to launch on another rocket, such as the Russian Soyuz-Fregat booster. Even as a secondary payload on a Soyuz mission, the launch capability of the booster enables a simpler design of the Cosmos 2 spacecraft compared with the design of Cosmos 1, therefore improving the probability of successful deployment of the solar sail upon reaching orbit. Aerocapture Aerocapture ranges from being a strongly enhancing to an enabling technology for the Neptune Orbiter with Probes and Titan Explorer missions and has been identified as an enhancing technology for robotic exploration of Mars and enabling for human missions to Mars. It is an orbit insertion maneuver that takes advantage of a planetâs atmosphere to decelerate a spacecraft sufficiently to allow it to be placed into its intended orbit. This type of maneuver is intended to minimize the use of propellants. It begins with the spacecraft entering the atmosphere of a target planet or moon. It then uses the friction or drag of the atmosphere to slow the vehicle. Once sufficient energy has been bled off, the spacecraft exits the atmosphere and performs orbit adjustment maneuvers to place it into the desired circular or elliptical orbit. Figure 3.6 depicts a generic aerocapture trajectory. The only propellant required for the overall maneuver is a minimal amount used for attitude control during atmospheric flight and for the final circularization maneuvers. Aerocapture requires guidance. Too shallow of an entry could result in the spacecraft skipping off the atmosphere and continuing into space permanently. Too steep of a trajectory could result in excessive aerodynamic and aero- thermal loads on the vehicle, potentially damaging or destroying it. It could also result in low apoapsis or impact with the target body (i.e., a crash). The guidance system, therefore, has to account for entry target and navigation errors, atmospheric uncertainties and variabilities, as well as aerodynamic uncertainties. Aerocapture guidance is analogous to that planned for use by Mars Science Laboratory and Orion to accomplish precision landing. â Masciarelli, Ball Aerospace and Technologies Corporation, âAerocapture Overview,â presentation to the Committee on Science Op- J. portunities Enabled by NASAâs Constellation System, June 9-11, 2008. â Rae Mayer, âIn-Space Propulsion Technology Program Office (ISPT) Overview,â February 2006.
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 77 Atmospheric Entry Interface Hyperbolic Approach Trajectory Begin Bank Angle Modulation Periapsis Circularization Periapsis Maneuver Raise Maneuver End Bank Angle Modulation Science Orbit Atmosphere Exit FIGURE 3.6â Aerocapture trajectory. SOURCE: Courtesy of Rae Mayer, In-Space Propulsion Technology Program Office, NASA. Figure 3.6.eps There are two basic types of aerocapture systems: aeroshells and deployables. The aeroshell designs are the most mature, having been used extensively for atmospheric entry missions. The deployable designs offer the benefit of lighter weight, but are much less mature. The aeroshell design encases the spacecraft in a structure covered with a thermal protection material. The shell serves as an aerodynamic surface, providing both lift and drag as the spacecraft flies through the target atmosphere. The thermal protection materials protect the structure from the intense aerodynamic and radiative heating experienced during atmospheric flight. The shell is then jettisoned once the spacecraft achieves its final orbit. There are blunt- and slender-body aeroshell configurations, with the blunt-body configuration being the most mature (see Figure 3.7). Blunt-body designs have been used extensively for atmospheric entry missions, such as the Mercury, Gemini, and Apollo crewed missions, as well as the Viking, Pioneer, Galileo, and the more recent Mars Exploration Rover (Sprit and Opportunity) missions. The slender-body design is less mature (Figure 3.8). However, it provides greater control authority as a lifting body and is more tolerant of navigation and atmospheric uncertainties. Studies have demonstrated that the slender- body design may be required for the Neptune mission. There are also two general families of deployable designs: trailing ballutes and inflatable aeroshells (Figures 3.9 and 3.10). A âballuteâ is an aerodynamic decelerator that is towed behind the vehicle. Inflatable aeroshells consist of a rigid shell with an attached inflated or deployed structure that increases the effective drag area. Bal- lutes rely strictly on aerodynamic drag for aerocapture, whereas the inflatable aeroshells also use lift for greater control during the atmospheric flight. Both types of deployable designs have the benefit of a drag area that can be significantly greater than the cross-sectional area of the spacecraft. As a result, they are effective at higher â NASA, Aerocapture TechnologyâIn-Space Propulsion Technology Project, NASAfacts FS-2007-09-12-GRC Pub ACap001, NASA Glenn Research Center, Cleveland, Ohio.
78 LAUNCHING SCIENCE FIGURE 3.7â Mars Science Laboratory aeroshell, which has a blunt-body shape similar to that required for aerocapture mis- sions at Mars and Titan. SOURCE: Courtesy of NASA.Figure 3.7.eps Bitmap image FIGURE 3.8â A potential slender-body aeroshield with atmospheric probes on top. This is the type of shape required for aero- capture at Neptune. SOURCE: Courtesy of NASA.
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 79 FIGURE 3.9â âBalloon-parachuteâ or âballute.â SOURCE: Courtesy of NASA. FIGURE 3.10â Inflatable aeroshell. SOURCE: Courtesy of NASA. altitudes, where the atmosphere is less dense and the aerodynamic and thermal loading are significantly lower. They can therefore be built with lower-mass materials and require less thermal protection, allowing more mass allocation for scientific payloads. A variant of aerocapture is skip entry. In this mode, the atmospheric entry trajectory is such that the vehicle dips into the atmosphere multiple times, decreasing the velocity of the vehicle with each dip. This reduces the overall heat load and aerodynamic loads associated with steeper descents. It also allows the targeting of a larger number of landing sites, or conversely allows the targeting of a given landing site from a wide range of entry points. This capability was developed for the Apollo program to allow the astronauts to target alternate Earth landing sites in the event of adverse weather conditions, but it was never used. The Russians used skip-entry trajectories for the return to Earth of both Zond 6 and Zond 7, two lunar flyby probes. The implementation of aerocapture requires an understanding of the target atmosphere through which the spacecraft will fly, as well of the aerodynamics and aerothermal environment of the spacecraft in the atmosphere. The design of the various spacecraft systems, including the guidance, navigation, and control (GN&C) system and associated algorithms, thermal protection system, and structures must accommodate the associated environments and flight conditions. A summary of the readiness of these aerocapture technologies is provided in Figure 3.11 for several space exploration missions. With the exception of Neptune missions, the technology is relatively mature, â TiborKremic, NASA, âNASA Investments in In-Space Propulsion Technologies,â presentation to the Committee on Science Opportunities Enabled by NASAâs Constellation System, June 9-11, 2008.
80 LAUNCHING SCIENCE Destination Venus Earth Mars Titan Neptune Subsystem Venus-GRAM based on Earth-GRAM validated by Mars-GRAM continuously Titan-GRAM based on Yelle Neptune-GRAM Atmosphere world-wide VIRA. Space Shuttle updated with latest atmosp. Accepted worldwide developed from Voyager, mission data. and updated with Cassini- other observations Goal: Capture Physics Huygens data Aerodynamics Heritage shape, well Heritage shape, well Heritage shape, well Heritage shape, well New shape; aerodynamics understood aerodynamics understood aerodynamics understood aerodynamics understood aerodynamics to be established. Goal: Errors = 2% GN&C APC algorithm captures Small delivery errors. APC APC algorithm captures APC algorithm captures APC algorithm with Î± Goal: Robust 96% of corridor algorithm captures 97% of 99% of corridor 98% of corridor control captures 95% of performance for 4-6 corridor corridor. DOF simulations More testing needed on Technology ready for ST9. ISPT investments have ISPT investments have Zoned approach for mass TPS efficient mid-density TPS. LMA hot structure ready for provided more materials provided more materials efficiency. Needs more Goal: Reduce SOA by Combined convective arrivals up to 10.5 km/s. ready for application to ready for application. investment. 30%+, expand TPS and radiative facility slow arrivals, and new choices needed. ones for faster entries. High-temp systems will High-temp systems will High-temp systems will High-temp systems will Complex shape, large Structures reduce mass by 31%. reduce mass by 14%-30%. reduce mass by 14%- reduce mass by 14%-30%. scale. Extraction difficult. Goal: Reduce SOA 30%. mass by 25% Convective models Environment fairly well- Convective models agree Convective models agree Conditions cannot be Aerothermal match within 20% known from Apollo, Shuttle. within 15%. Radiative: within 15%. Radiative duplicated on Earth in laminar, 45% with Models match within 15% predict models will agree models agree within 35- existing facilities. More Goal: Models match turbulence. Radiative within 50% where radiation 300% work on models needed. within 15% models agree within 50% is a factor. System Accomplishes 97.7% of Accomplishes 97.2% of âV Accomplishes 97.8% of Accomplishes 95.8% of Accomplishes 96.9% of Goal: Robust âV to achieve 300 x 300 to achieve 300 x 130 km âV to achieve 1400 x 165 âV to achieve 1700 x 1700 âV to achieve Triton performance with km orbit. No known orbit. No known technology km orbit. No known km orbit. No known tech observer. orbit. ready technology technology gaps. gaps. technology gaps. gaps. ENABLING ENABLING Ready for Infusion Some Investment Needed Significant Investment Needed FIGURE 3.11â Summary of aerocapture technology readiness for several different targets. Whereas aerocapture could be e Â asily developed for use at Venus, Earth, Mars, and Titan, developing it for use at Neptune will require significant investment Figure 3.14.eps in several important areas (note that the chart indicates that extracting a spacecraft from its aeroshell will be difficult for the Neptune mission). According to this NASA source, aerocapture is required for Titan and Neptune missions and is therefore âenabling.â An Ares V launch vehicle could potentially alter that conclusion. NOTE: GRAM, Global Reference Atmospheric Model; VIRA, Venus International Reference Atmosphere; GN&C, guidance, navigation, and control; DOF, degrees of free- dom; APC, Analytic Predictor Corrector; TPS, thermal protection system; SOA, state of the art; ST9, Space Technology 9; LMA, Lockheed Martin Astronautics; ISPT, In-Space Propulsion Technology. SOURCE: Courtesy of NASA. but additional development work is necessary. The unique environment of Neptune will require significant addi- tional development effort to enable exploration. Significant maturity has been gained in the areas of atmospheric modeling, aerodynamics, and GN&C. The atmospheric composition and models have been well established and accepted by the community for all destina- tions. A simple blunt-body aeroshell design can be used for the majority of the destinations. These systems have a long heritage, dating back to the early crewed space program, and the associated aerodynamics have been well characterized. Missions to Neptune, however, may require a slender-body aeroshell. There has been less work in this area, so additional aerodynamic characterization is required. A robust guidance algorithm, the Analytic Predictor Corrector, has been developed and shown through simulation to be able to fly successfully through atmospheres of any of the destinations and deliver the spacecraft to the desired orbit. This algorithm is able to accommodate the uncertainties and variabilities of the various flight parameters, including atmospheric variations and perturbations and uncertainties in the aerodynamic conditions and spacecraft mass properties, as well as the navigation errors that build up in flight. Currently NASA has no plans to develop or demonstrate aerocapture beyond its current TRL and has recently determined that this technology will not be developed further. This decision will severely limit the science explo-
TECHNOLOGY REQUIREMENTS FOR FUTURE SPACE MISSIONS 81 ration of Neptune even with an Ares V capability (for instance, by preventing the mission from carrying both atmospheric probes and a Triton lander) and could have an adverse impact on missions to Venus and Titan. PROPULSION SYSTEM TECHNOLOGY SUMMARY In the recent past, NASA undertook a number of impressive in-space propulsion technology development projects that reached moderate to high technology readiness levels and demonstrated significant promise for future missions. However, for various reasons the agency has eliminated much of this research. The committee concluded that at least some of these technologies will be required for future missions even if they use the capabilities of the Constellation System. Finding: Advanced in-space propulsion technology may be required for some science missions considered for using the Constellation System. THE DEEP SPACE NETWORK The Deep Space Network, developed during the Apollo era, consists of three major tracking sites spaced in longitude around the globe (in Spain, California, and Australia) to provide continuous communication and naviga- tion support for deep space (beyond geosynchronous) missions. (See Box 3.2.) Many of the mission concepts that were evaluated by the committee would place additional demands on the DSN (see Table 3.1). These additional demands are likely to arise from several sources: â¢ The mission concepts are large, complex payloads with concomitant large outputs of data. For example, large-aperture telescopes, operating at the Sun-Earth L2 point with gigapixel focal planes, would produce a sub- stantial amount of data at a relatively far distance from Earth. While precise definition of data transmission and uplink requirements are stated only in some of the mission concepts, it is clear that during the time period of this study (2020-2035), data transmission requirements will grow substantially. â¢ A number of mission concepts envision servicing of the scientific payloadsâeither robotically or by way of human visits. This implies a need for substantial two-way communications to Earth uplinks and downlinks and increased data flow. Assured connectivity and link redundancy will be required if human repair and servicing activities on orbit are involved. In addition, a mission to send astronauts to a Near Earth Object would require substantial two-way communications for the same reasons. â¢ Some of the concepts that the committee evaluated would stress the capabilities of DSN simply because of the very weak signals at the communication distances involved. Concepts of orbiters around Neptune at 30 AU or a mission to the boundary of interstellar space to about 200 AU are two examples of types of missions for which large-aperture antennas on Earth would be required. The committee requested and received information from NASA about the planned course of evolution of the Deep Space Network. It learned that by 2025 the DSN may not reliably meet the current and projected uplink and downlink requirements. These current and projected requirements do not include any of the Constellation- enabled missions reviewed during this study, which nominally would fly in the 2020-2035 time interval. NASA is concerned with the deterioration of the existing infrastructure, particularly the 70-m antennas now more than 40 years old, and the lack of overall investment to modernize the DSN to accommodate the expected increase in data rates from a variety of missions, including missions to the Moon and Mars. Absent such modernization, the data return from the Constellation-enabled missions might be compromised and fail to fulfill the potential of these very expensive missions. Finding:â Science missions enabled by the Constellation System will increase the strain on the capabilities of the Deep Space Network.
82 LAUNCHING SCIENCE BOX 3.2 Brief Description of the Deep Space Network The Deep Space Network (DSN) is an international network of antennas that supports interplan- etary spacecraft missions, selected Earth-orbiting missions, and radio and radar astronomy observations for the exploration of the solar system and the universe. The network is a NASA facility managed by the Interplanetary Network Directorate (IND) within the Jet Propulsion Laboratory (JPL). (See Figure 3.2.1.) The origins of the DSN can be traced back to the late 1950s and the inception of NASA in 1958. The execution of several planetary robotic missions, managed by JPL, also contributed to the further develop- ment of the network, which expanded from a single tracking station to a worldwide network of large-dish antennas. The DSN consists of three deep space communications facilities strategically placed approxi- mately 120 degrees from one another. The Goldstone Deep Space Communications Complex is located in Californiaâs Mojave Desert; the Madrid Deep Space Communications Complex is in Spain, 37 miles to the west of Madrid at Robledo de Chavela; and the Canberra Deep Space Communications Complex is in the Australian Capital Territory, 25 miles southwest of Canberra near the Tidbinbilla Nature Reserve. Each complex is situated in semimountainous, bowl-shaped terrain for better shielding against radio-frequency interference, and consists of at least four deep space stations equipped with ultrasensi- tive receiving systems and large parabolic dish antennas. The system is made up of the following array of antennas: one 34-meter-diameter high-efficiency antenna, one or more 34-meter beam waveguide anten- nas (three at Goldstone, two in Madrid, and one in Canberra), one 26-meter antenna, and one 70-meter antenna. Every one of these is a steerable, high-gain, parabolic reflector antenna. All of the stations are remotely oper- ated from a centralized Signal Processing Center at each complex, which house the electronic subsystems that control the anten- nas, receive and process the data, transmit commands, and generate the spacecraft navi- gation data. The data are then transmitted to JPL for further processing and distribution to science teams over a ground communications network. The antennas and data delivery sys- tems make it possible to acquire telemetry data and transmit commands to spacecraft, track spacecraft position and velocity, perform very long baseline interferometry observa- tions, measure variations in radio waves for radio science experiments, gather science data, and monitor and control the perfor- mance of the network. The strategic placement of these facili- ties is what characterizes the DSN as the larg- FIGURE 3.2.1â One of the 70-meter dishes of the Deep Space est and most sensitive scientific telecommu- Network. This system currently has long-term maintenance nications system in the world. Their position and usage challenges, but it would be significantly taxed by many of the proposed mission concepts evaluated in this permits the constant observation of spacecraft study. SOURCE: Courtesy of NASA. and enables continuous observation and suit- able overlap for transferring the spacecraft radio link from one complex to the next.