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Optimization of extended propulsion time nuclear-electric propulsion trajectories
This paper presents the methodology used in optimizing extended propulsion time NEP missions considering realistic thruster lifetime constraints. These missions consist of a powered spiral escape from a 700-km circular orbit at the earth, followed by a powered heliocentric transfer with an optimized coast phase, and terminating in a spiral capture phase at the target planet. This analysis is most applicable to those missions with very high energy requirements such as outer planet orbiter missions or sample return missions where the total propulsion time could greatly exceed the expected lifetime of an individual thruster. This methodology has been applied to the investigation of NEP missions to the outer planets where examples are presented of both constrained and optimized trajectories.
Benefits of Nuclear Electric Propulsion for Outer Planet Exploration
Nuclear electric propulsion (NEP) offers significant benefits to missions for outer planet exploration. Reaching outer planet destinations, especially beyond Jupiter, is a struggle against time and distance. For relatively near missions, such as a Europa lander, conventional chemical propulsion and NEP offer similar performance and capabilities. For challenging missions such as a Pluto orbiter, neither chemical nor solar electric propulsion are capable while NEP offers acceptable performance. Three missions are compared in this paper: Europa lander, Pluto orbiter, and Titan sample return, illustrating how performance of conventional and advanced propulsion systems vary with increasing difficulty. The paper presents parametric trajectory performance data for NEP. Preliminary mass/performance estimates are provided for a Europa lander and a Titan sample return system, to derive net payloads for NEP. The NEP system delivers payloads and ascent/descent spacecraft to orbit around the target body, and for sample return, delivers the sample carrier system from Titan orbit to an Earth transfer trajectory. A representative scientific payload 500 kg was assumed, typical for a robotic mission. The resulting NEP systems are 100-kWe class, with specific impulse from 6000 to 9000 seconds.
Dual-Telescope Multi-Channel Thermal-Infrared Radiometer for Outer Planet Fly-By Missions
The design of a versatile dual-telescope thermal-infrared radiometer spanning the spectral wavelength range 8-200 microns, in five spectral pass bands, for outer planet fly-by missions is described. The dual- telescope design switches between a narrow-field-of-view and a wide-field-of-view to provide optimal spatial resolution images within a range of spacecraft encounters to the target. The switchable dual-field- of-view system uses an optical configuration based on the axial rotation of a source-select mirror along the optical axis. The optical design, spectral performance, radiometric accuracy, and retrieval estimates of the instrument are discussed. This is followed by an assessment of the surface coverage performance at various spatial resolutions by using the planned NASA Europa Mission 13-F7 fly-by trajectories as a case study.
Magnetometer instrument team studies for the definition phase of the outer planets grand tour
The objectives of magnetic field investigations on missions to the outer planets were defined as well as an instrumentation system, a program of studies and instrument development tasks was proposed for the mission definition phase of the Outer Planets Grand Tour project. A report on the status of this program is given. Requirements were also established for the spacecraft and the mission which would insure their compatibility with the magnetic field investigation proposed for the outer planets missions and developed figures of merit for encounter trajectories. The spacecraft-instrumentation interface and the on-board data handling system were defined in various reports by the Project Team and in the reports by the Science Steering Group. The defining program for exploring the outer planets within the more restrictive constraints of the Mariner Jupiter-Saturn project included defining a limited magnetic field investigation.
Data subsystems for 12-year missions
Thermoelectric outer planet spacecraft /TOPS/ data subsystems for 12 year missions compared with Mariner subsystem
Satellite encounter opportunities of the Grand Tour missions
Multiple outer planets flyby opportunities during 1976-1980 Grand Tour missions launch period, presenting computer generated specific encounter dates and flyby distances
Ultrasail
Ultrasail is a complete sail system for the launch, deployment, stabilization and control of very large solar sails enabling reduced mission times for interplanetary and deep space spacecraft. Ultrasail is an innovative, non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation-flying microsatellites with an innovative solar sail architecture to achieve sq km-class controllable sail areas, sail subsystem area densities of 1 gm per sq m, and thrust levels equivalent to 400 kW ion thruster systems used for comparable deep space missions. Ultrasail can conceivably even achieve outer planetary rendezvous, a deep space capability now reserved for high-mass nuclear and chemical systems. Ultrasail is a Delta IV-launched multi-blade spin-stabilized system with blade lengths as long as 50 km, reminiscent of the MacNeal Heliogyro. The primary innovation is the near-elimination of sail supporting structures by attaching the sail tip to a rigid formation-flying microsatellite truss which deploys the sail blade, and which then articulates the blade to provide attitude control, including spin stabilization and precession of the spin axis. These tip microsatellites are controlled by a solar-powered 3-axis microthruster system (electric or cold gas) to maintain proper sail film tension during deployment and spin-up. The satellite mass also provides a stabilizing centrifugal force on the blade while in rotation. Understanding the dynamics of individual blades is key to the overall dynamics of Ultrasail. Forces and torques that must be modeled include those due to solar pressure, those generated by the microsatellite at the blade tip and by torques applied at the blade root. Centrifugal forces also play a significant role in the deployment and maintenance of the sail configuration. To capture the dynamics of the overall system, the equations of motion for the blades have been derived. Using these differential equations, a control law will be derived to maneuver Ultrasail. This law involves the pitching of the individual blades thereby moving the distribution of the radiation pressure on each individual blade and inducing a resultant torque on the system. The direction of the angular momentum vector and its rate of precession can be controlled through the pitch angle of the blades. The Ultrasail trajectory is also being studied. Optimal or near-optimal trajectories are being generated to showcase Ultrasail performance. Various missions, e.g. outer planet and solar polar missions for observation of the Sun, are currently being investigated to demonstrate the performance enhancements generated by Ultrasail technology. Calculus-of-variations-based optimization software is used to produce optimal Ultrasail trajectories. The performance of these trajectories is being compared to optimal results generated with other propulsion models, including chemical propulsion, ion propulsion, and competing solar sail concepts. Results of these studies will quantify the performance of Ultrasail compared to existing solar sail concepts for high energy missions.
Parameterization of Nuclear Electric Propulsion Trajectories for Outer Solar System Science Missions Methodology
This manuscript addresses the methodology used to create a database of low thrust missions to outer planets. This database utilizes previous work modeling NEP systems to determine the maximum delivered mass to outer planets based on a range of mission parameters, such as time of flight, launch vehicle, and power system mass. Trajectories were selected which best utilized NEP benefits. Additionally, a discussion on the database outputs for missions to Saturn is included, such as time of flight based on trajectory type and maximum payload, given a specific launch vehicle. The purpose of this work was to create a basis for future mission design, and a tool to investigate general trends across mission options.
Lithium-fuelled electromagnetic thrusters for robotic and human exploration missions
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Solar electric low thrust missions to Jupiter with swingby continuation to the outer planets.
Solar electric low thrust propulsion application to Jupiter swingby mission to outer planets
Solar electric low-thrust missions to Jupiter with swingby continuation to the outer planets.
Solar electric low thrust propulsion application to Jupiter swingby mission to outer planets
Solar electric propulsion mission and spacecraft capabilities for outer planet exploration
This paper provides an update on solar electric-propulsion (SEP) technology by describing how SEP, as currently understood, performs relative to chemical-ballistic options on a limited set of missions for exploring the outer planets. A spacecraft design for these missions is discussed which embodies the findings of a study and development effort by NASA's Office of Space Science and Office of Advanced Systems Technology.
Next Generation Thermal Protection System for Outer Planet Probes and Orbiter
Saturn Probe and Ice Giant Orbiter along with in-situ Probe Missions continue to be very high priority mis-sions. A result of the advocacy by OPAG and other Analysis Groups, the 3D Woven, Dual-Layer HEEET thermal protection system, mature at TRL 6, has closed the TPS gap for extreme environment missions. A mid density follow on to DL HEEET, developed to meet earth entry requirements, is a Single Layer vari-ant of the 3D woven TPS, which provides a mass effi-cient single layer 3D Mid-Density TPS (3MDCP) that has been baselined as the heatshield for MSR EEV. Continued development of 3MDCP will elevate it to TRL 6 by 2025. Conformal-PICA (C-PICA) devel-opment was pursued to establish a more efficient and robust alternate to PICA, and it is at TRL 4+. Our rationale for the next generation of TPS de-velopment is based on the on the missions needs of the next decade Outer Planet missions that are unique and more demanding than any other destinations. Taking advantage of the recent planned development of sev-eral materials, the next generation of TPS offers a much more mass efficient option for small, medium, and large class Outer Planet missions. Saturn Probes: While mission designers are gen-erally interested in shallow entry to maintain the g-load during entry around 50g, the heat-load for shallow entry can range between (100 kJ/cm2 – 300 kJ/cm2), two orders of magnitude higher than Venus or Sample Return missions. TPS must not only offer protection but must be mass efficient to perform reasonable sci-ence. TPS mass can quickly become 50% or more of the mass of the entire entry system. Recent analysis performed shows Saturn Probe missions could signifi-cantly benefit from the single layer HEEET (SL-HEEET/3MDCP). A DL-HEEET based heatshield mass could be ~ (40% -50%), SL-HEEET can provide additional (30% - 50%) mass savings. At the same time, C-PICA can provide (30% - 50%) mass savings on the backshell. The combined mass savings can be significant enough to carry an additional probe, if de-sired. Ice Giant Aerocapture Missions: Aerocapture mission architectures can provide significant ad-vantage over traditional propulsive insertion missions in multiple ways. 1) Reduced trip time ~ (4- 6) years (30% -40%), 2) Enables placing the orbiter, probe, and lander, all together and 3) Allows for greater science mass (probes and landers) due to mass efficiency. The delivery of a probe from orbit makes it easier and eliminates mission design constraints by HEEET for direct entry [4] of probes and allows for more targeted in-situ science once the Orbiter is able to collect data. In the past 20 years, progress made in GN&C for lift-guided entry missions (MSL, Orion EFT1, Mars 2020) and the expertise in blunt body aerodynamics at large scale (~ 5m) has led the EDL community to conclude that aerocapture is a “go do” engineering activity. An aerocapture mission that will deplete the excess energy of a fast arrival mission will require a mass efficient TPS that can handle extreme heat-load, ~ (100 kJ/cm2 – 500 kJ/cm2). Hence TPS, feasibility as well as mass efficiency requires assessment. Utilizing the recent developments, a comprehensive, bounding analysis was done to establish the potential for SOA (HEEET) system as well as emerging new TPS such as SL-HEEET and C-PICA. In this proposed poster, we will outline the process by which we establish bounding aerocapture trajecto-ries for hyperbolic excess velocities ranging from 27 km/s to 35 km/s, for low L/D (~ 0.4) configurations and determine conservative/bounding estimate of aer-othermal environment by using a combination of CFD simulations and stagnation point heating estimates [7]. This engineering approach allows us to assess the TPS need vs. TPS capability and determine the applicability of existing TPS. Once an applicable suite of TPS is determined, the TPS thickness and mass are comput-ed. We show that the TPS mass fraction can be as low as 5% to as high as 20%, depending on the use of ad-vanced TPS, while HEEET is sufficient but will require 50% of the entry mass.
Aerocapture Technology Development Needs for Outer Planet Exploration
The purpose of this white paper is to identify aerocapture technology and system level development needs to enable NASA future mission planning to support Outer Planet Exploration. Aerocapture is a flight maneuver that takes place at very high speeds within a planet's atmosphere that provides a change in velocity using aerodynamic forces (in contrast to propulsive thrust) for orbit insertion. Aerocapture is very much a system level technology where individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), guidance, navigation and control (GN&C) instrumentation need to be integrated and optimized to meet mission specific requirements. This paper identifies on-going activities, their relevance and potential benefit to outer planet aerocapture that include New Millennium ST7 Aerocapture concept definition study, Mars Exploration Program aeroassist project level support, and FY01 Aeroassist In-Space Guideline tasks. The challenges of performing aerocapture for outer planet missions such as Titan Explorer or Neptune Orbiter require investments to advance the technology readiness of the aerocapture technology disciplines for the unique application of outer planet aerocapture. This white paper will identify critical technology gaps (with emphasis on aeroshell concepts) and strategies for advancement.
Optimal lunar trajectories for a combined chemical-electric propulsion spacecraft
Spacecraft which utilize electric propulsion (EP) systems are capable of delivering a greater payload fraction compared to spacecraft using conventional chemical propulsion systems. Several researchers have investigated numerous applications of low-thrust EP including a manned Mars mission, scientific missions to the outer planets, and lunar missions. In contrast, the study of optimal combined high and low-thrust spacecraft trajectories has been limited. In response to the release of NASA's 1994 Announcement of Opportunity (AO) for Discovery class interplanetary exploration missions, a preliminary investigation of a lunar comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed. The Discovery mission (eventually named Diana) was envisioned to be a two-phase scientific exploration mission: the first phase involved exploration of the moon and second phase involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low-thrust SEP transfer to a circular, polar, low-lunar orbit (LLO). After scientific data was collected at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a short period comet. An initial study that demonstrated the feasibility of using EP for the lunar and comet orbit transfer was performed under the grant NAG3-1581. This final report is a continuation of the initial research efforts in support of the Discovery mission proposal that was submitted to NASA Headquarters in October 1994. Section 2 discusses the lunar orbit transfer phase of the Diana mission which involves both chemical and electric propulsion stages. Section 3 discusses the chemical lunar orbit insertion (LOI) burn optimization. Finally, section 4 presents the conclusions of this research effort.
LO2/LH2 propulsion for outer planet orbiter spacecraft
Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.