Interstellar explorerjun01
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This document discusses a proposed mission concept to send a probe to 1000 AU within 50 years using currently feasible technologies. Key elements include: 1) Using a solar gravity assist at Jupiter to eliminate angular momentum, then falling to within 4 solar radii of the Sun to leverage the high speeds for an escape trajectory. 2) The probe would use a high-Isp propulsion system like solar thermal or nuclear thermal to accelerate during a 15-minute perihelion maneuver for solar system escape. 3) Enabling technologies discussed include high-temperature carbon-carbon shields, efficient radioisotope power, and laser optical communications. Follow-on studies are proposed to further develop the concepts.
![[Showed Epsilon Eridani Movies]](https://image.slidesharecdn.com/interstellarexplorerjun01-140902075833-phpapp01/85/Interstellar-explorerjun01-12-320.jpg)
![Thermal Propulsion
Chemical propulsion cannot provide sufficiently high Isp due to the
high mean molecular weight of the combustion products
Suggests using low-molecular weight and a decoupled energy
source
Solar Thermal Propulsion (STP) - tap the Sun’s energy
via the thermal shield - Being studied here
Nuclear Thermal Propulsion (NTP) - use a compact ultra-low
mass MInature ReacTor EnginE (MITEE) [Powell et
al., 1999]
Advanced architecture (U-233 fuel, BeH2 moderator, LH2 reflector )
could provide criticality in a ~40 kg package.
Further decrease could come from Am-242m fuel (supply issues !!)
Question of possible size of Pu-239 system (plenty of fuel!)
Needs further study at the systems level](https://image.slidesharecdn.com/interstellarexplorerjun01-140902075833-phpapp01/85/Interstellar-explorerjun01-14-320.jpg)
![Schedule
2000-2002 Advanced Technology Development study(ies)
2000-2002 Continued definition studies of the solar sail concept for IP at JPL
2002-2003 Update of OSS strategic plan with study for a "New Millennium"-like mission
2003-2007 Focused technology development for small probe technologies
2004-2007 Development of sail demonstration mission
2004-2007 Development of Solar Probe mission (test for perihelion propulsion)
[2006-2007 Hardware tests for radioisotope sail feasibility ]
[2006-2007 Hardware tests for antimatter propulsion schemes ]
2006-2007 Monitor DoD STP effort and conduct NASA-specific hardware tests
[2002-2007 Development of space-qualified nuclear thermal reactor ]
2007-2010 Focused technology development for an Interstellar Probe
2009-2012 Design and launch of first generation solar-sail probe
2010 Test of Solar Probe performance in the perihelion pass of October 2010
2012-2015 Design and launch second generation probe 1000 AU goal in 50 years
2015-2065 Data return from 1000 AU and “beyond the infinite...”](https://image.slidesharecdn.com/interstellarexplorerjun01-140902075833-phpapp01/85/Interstellar-explorerjun01-36-320.jpg)
Interstellar explorerjun01
- 1. 1 Ralph L. McNutt, Jr. The Johns Hopkins University Applied Physics Laboratory Laurel, MD 20723-6099 June 5, 2001 NASA Institute for Advanced Concepts 3rd Annual Meeting NASA Ames Research Center
- 2. Other Contributors G. Bruce Andrews System Engineering Robert S. Bokulic RF Communications Bradley G. Boone Optical Communications David R. Haley Guidance and Control J. V. McAdams Mission Design M. E. Fraeman Ultra-Low Power Electronics B. D. Williams Thermal Design M. P. Boyle Mechanical Design D. Lester, R. Lyman, M. Ewing, R. Krishnan - Thiokol Corp D. Read, L. Naes - Lockheed-Martin ATC M. McPherson, R. Deters - Ball Aerospace Support From: Task 7600-039 from the NASA Institute for Advanced Concepts (NIAC) under NASA Contract NAS5-98051
- 3. Starship Concepts Have Intrinsically Large Dry Masses (100s of Tons) Technology Extrapolations Sound “Too Good To Be True” - and May Be ... Driven by propulsion requirements Driving costs to scale of current GDP Daedalus Fusion Rocket (D-3He) Sänger Photon Rocket Bussard Ramjet
- 4. A Mission to the VLISM is More Modest - But Can be Done in the Near Term The external shock may be ~300 AU away - so 1000 AU is “clear”of the influcence of the Sun on its surroundings
- 5. Mission Concept Reach a significant penetration into the Very Local Interstellar Medium - out to ~1000 AU - within the working lifetime of the probe developers (<50 years) To reach high escape speed use a solar gravity assist (due to Oberth, 1929): (1) Launch to Jupiter and use a retrograde trajectory to eliminate heliocentric angular momentum (2) Fall in to 4 solar radii from the center of the Sun at perihelion (3) Use an advanced-propulsion system ΔV maneuver to increase probe energy when its speed is highest to leverage rapid solar system escape
- 6. Enabling Technologies High Isp, high-thrust propulsion (for perihelion maneuver, ~15 minutes) Carbon-carbon thermal shield Long-range, low-mass telecommunications Efficient Radioisotope Thermoelectric Generator (RTG) Low-temperature (<150K), long-lived (<50yr) electronics <0.1 arc second pointing for data downlink Open loop control Fully autonomous operational capability with onboard fault detection and correction Possible extension to multi-century flight times while maintaining data taking and downlink operations
- 7. Phase I Study Topics Architectures that allows launch on a Delta III-class vehicle Redundancies that extend probe lifetime to >1000 years; software autonomy, safing Concept that links science, instruments, spacecraft engineering, and reality 1000 AU, 50-year mission with extension to 1,000 years (~20,000 AU) Optical downlink to support 500 bps at 1000 AU Propulsion concepts, e.g., Solar Thermal, Nuclear Pulse, Nuclear Thermal - to enable the perihelion burn
- 8. Phase II Effort Proposed to Refine and Develop Phase I Concepts Refine development of consistent thermal, propulsion, and mechanical design - In process Examine use of transuranic isotopes in propulsion and power system - Done Conduct STP and NTP system designs and trades including propellant selection and storage - Done Breadboard and program self-healing, distributed-processor spacecraft architecture to demonstrate use and resiliency - In process Develop optical-communication concept - Maturing Examine trades against low-thrust propulsion concepts - Starting
- 9. Approach is to Maximize Use of Talent Base Goal is the development of a “realistic” concept Needs dreams as well as solid engineering approach So .... Solicit input from lead engineers delivering flight hardware - who also have proven track records Best solution is find personnel who have ALSO worked TRL 4-6 ATD programs Use NIAC funds (limited) to maximize technical work by “fitting in” around ongoing flight and ATD programs (ensures best engineering talent - which is also typically oversubscribed)
- 10. Mission Design Solar system escape speed is set by the ΔV at perihelion and the radial distance from the center of the Sun rp v V 35.147 r escape p = (Δ ) 1 2 1 4 1 AU/yr = 4.74 km/s A "now-technology" Interstellar Probe could supply a ΔV of 1.56 km s-1 at perihelion, about one-tenth of what is desirable. Perihelion distance = 3 RS => ~7.0 AU yr-1 To reach ~20 AU yr-1 the probe needs to be accelerated by ~10 to 15 km s-1 during about 15 minutes around perihelion to minimize gravity losses
- 11. Trajectory toward ε Eridani Launching toward a star enables comparison of local properties of the interstellar medium with integrated properties determined by detailed measurements of the target-star spectrum, so we target the Sun-similar star ε Eridani, a K2V dwarf main sequence star 10.7 light years from Earth
- 12. [Showed Epsilon Eridani Movies]
- 13. Nuclear Pulse Propulsion Pulsed fission can, in principle, provide the key element for the perihelion propulsion. For a 260 kg probe (incld. 30% margin) and 215 kg of propellant, we need the fission energy of ~1.3 g of uranium - a total of about 13 tons of TNT equivalent The problem is the coupling of the momentum into the ship over short time scales ~10-8s and this is exacerbated by yields of ~1 to 10 kT explosions. Scalability to low-mass systems is problematic due to critical mass of fission assemblies Even then most promising known transuranic elements do not solve the problem - Np-236, Pu-241, Am-242m, Cm-245, Cf-249, and Cf-251 evaluated and compared with U-233, U-235, and Pu-239 Conclusion: Cannot be applied to “small” systems
- 14. Thermal Propulsion Chemical propulsion cannot provide sufficiently high Isp due to the high mean molecular weight of the combustion products Suggests using low-molecular weight and a decoupled energy source Solar Thermal Propulsion (STP) - tap the Sun’s energy via the thermal shield - Being studied here Nuclear Thermal Propulsion (NTP) - use a compact ultra-low mass MInature ReacTor EnginE (MITEE) [Powell et al., 1999] Advanced architecture (U-233 fuel, BeH2 moderator, LH2 reflector ) could provide criticality in a ~40 kg package. Further decrease could come from Am-242m fuel (supply issues !!) Question of possible size of Pu-239 system (plenty of fuel!) Needs further study at the systems level
- 15. Interstellar Probe Thermal Requirements • Survive cruise mode prior to perihelion pass Protect propellant system • Survive high heating rates at 4RS (2900 Suns) • Allow perihelion burn to accelerate vehicle • Deploy probe after burn • Use waste heat from RTG (or equivalent) to minimize heater-power requirements Operate probe electronics at ~ 125 K
- 16. Trade Studies Concentrate on STP system - results also apply to NTP Sufficiently large Isp to provide V Examine LH2, CH4, NH3 Maximize propellent temperature (up to structural failure) Examine pressure vs flow rate, heating, and recombination Size propellant tank/cryostat for propellant requirements Storage for cruise Pressure and expulsion during burn
- 17. Solar Thermal Propulsion Concept Pressu Temperature (K) re (kPa) 2400 3000 3300 3500 517 860 1037 1166 1267 H2 69 875 1144 1336 1369 † 517 480 588 667 705†† CH4 69 485 628 698† 705†† 517 421 502 559 604 NH3 69 427 547 634 639†† † † Pressure = 165 kPa †† Pressure = 910 kPa ††† Pressure = 221 kPa Hydrogen (H2) Specific Impulse (Vacuum ) Relative to Temperature and Nozzle Expansion Ratio Pchamber = 1380 kPa 1300 1200 1100 1000 900 800 700 600 20 30 40 50 60 70 80 90 100 Nozzle Area Expansion Ratio Isp vac (sec) 1500°K 1750°K 2000°K 2400°K 3000°K 3300°K 3500°K The baseline propellant hydrogen shows the most promise for obtaining the maximum ISP level Maximum achievable ISP with NH3 and CH4 are 639s and 705s, respectively (at 3500K)
- 18. Parametric Analysis • Incident heat flux: – 381 W/cm2 at 4 Rs – 396 W/cm2 at 3 Rs • # of plies: – 1 (0.3 mm) – 2 (0.6 mm) – 3 (0.9 mm) • Spacing, s – 5 mm – 10 mm – 15 mm • Mass flow rate – 200 g/s – 1100 g/s – 2000 g/s Test Case q"inc (W/cm2) # plies spacing (mm) mdot (g/s) pin (Pa) hx1 381 1 10 200 500000 hx2 381 3 5 200 500000 hx3 381 3 10 200 500000 hx4 381 3 15 200 500000 hx5 381 3 5 1100 1500000 hx6 381 3 10 1100 500000 hx7 381 3 15 1100 500000 hx8 381 3 5 2000 1800000 hx9 381 3 10 2000 1800000 hx10 381 3 15 2000 1800000
- 19. Results – hx2 hx2 3000 2500 2000 K) (1500 T 1000 500 0 x (m) T (K) Ts (K) 0 1 2 3 4 5 6 hx2 80 70 60 50 40 30 20 10 0 0 1 2 3 4 5 6 x (m) p (psi) p (psi)
- 20. Results – hx5 hx5 2500 2000 1500 1000 500 0 0 1 2 3 4 5 6 x (m) T (K) T (K) Ts (K) hx5 250 200 150 100 50 0 0 1 2 3 4 5 6 x (m) p (psi) p (psi)
- 21. Heat Shield Structural Evaluation • The 5mm and the 10mm cell size configurations were evaluated for structural integrity – Six different wall thicknesses (1ply to 6ply, 0.3mm each) – Typical 3D carbon-carbon material properties used @ 3000F – 200 psi fluid pressure assumed – Stress criteria used to determine acceptable configurations Maximum Stresses (Allowable stress ~ 14 ksi) CELL 1ply 2ply 3ply 4ply 5ply 6ply 5mm 23,740 4,278 1,996 1,249 874 658 10mm 91,890 22,540 6,571 4,044 2,779 1,930
- 22. Heat Shield Structural Evaluation Deformation (Typical) Stresses (Typical) Relative weight of a 1 inch specimen CELL (mm) # cells total area 2ply wt (lbs.) 3ply wt (lbs.) 5ply wt (lbs.) 5 1730 1021.589 1.56 2.33 3.89 10 868 1025.299 1.56 2.34 3.90 15 581 1029.01 1.57 2.35 3.92
- 23. LH2 Storage Options Best design is for low-pressure system with graphite epoxy; launch with solid LH2 and gradually melt prior to perihelion
- 24. Configuration Evolution Driven by LH2 Volume (1) Initial Concept (2) Maximize volume for Delta III (3) Size diven by 250 kg (dry) cryostat (4) Stack probe and cryostat shield in 5-m shroud
- 25. Thermal Constraints Are Met Analysis with Thermal Synthesis System (TSS) software CC primary shield with 0.85/0.55 at temperature (2964K) ~100 kg of CC aerogel backing on primary shield “Fins” on sides of cryostat capture energy to exhaust propellant
- 26. LH2 Thermally Isolated Until Needed Flat Plate Shield Concept for Instellar Probe 250 kg Cyrostat (No Radiator on Cryostat), Probe piggy back 3000 2500 2000 1500 1000 500 0 -500 -300 -200 -100 0 100 200 300 Time (Hours) Temperature (C) Shield Front Side Shield Back Side CRYOSTAT SIDES (MLI) CRYOSTAT WALL FWD CRYOSTAT WALL Flat Plate Shield Concept for Instellar Probe 250 kg Cyrostat (No Rad. on Cryostat), Probe piggy back 0 -50 -100 -150 -200 -250 -300 -300 -200 -100 0 100 200 300 Time (Hours) Temperature (C) PROBE INTERNAL PROBE SIDES (MLI) PROBE FWD END (MLI) Overall temperatures Internal temperatures
- 27. Current Concept Accomodates 400 kg LH2 Protect LH2 with thermal shield Keep CG in line with thrust Propellant lines connect tank to shield and to DeLaval nozzle and to perihleion heat exchangers Fits in 5-m shroud Primary and secondary thermal shields Adaptor ring shown Probe rides in shadow of propellant tank
- 28. Interstellar Probe Final Flight Configuration 50 kg, 15 W probe Operate at ~125K Includes: 10 kg, 10W Science Instruments Following the perihelion burn, the probe consists of three main mechanical elements – an RPS, a central support mast containing the comm laser and battery, and an optical dish pointing toward the solar system. Instruments and processors mount to the back of the dish.
- 29. 00-0642-15 Probe Block Diagram Perihelion propulsion module is not shown
- 30. 00-0642-1 Wireless Communication Module • 2.4 GigaHz operation • 72 channels • Two will be interfaced to each ultra low power processor using the serial RS-232 ports (which support 56 K-baud communication) • One is used for inter-processor communication, the other for communication with subsystems
- 31. 00-0642-1 Typical S/C Architecture (MESSENGER)
- 32. 00-0642-2 Why are Simple Dual Redundant Systems the Current “Standard”? • Good flight reliability history for missions < 10 years long. Why change? • Ultra low power (ULP) processors, which would enable more redundancy on a S/C, are not flight ready. • Even if ULP processors where available now, cross-strapping S/C subsystems between >4 processors is cumbersome. • RF links for inter-processor communication, as well as with S/C subsystems and instruments, enable n-way cross-strapping, but they, too, are not flight-worthy at this point in time
- 33. 00-0642-3 Advantages of Interstellar Probe S/C Architecture over Current “Standard” • Each ULP processor on IP is powerful enough to run S/C operations by itself, so if IP has “N” processors then IP has true “N”-fold redundancy – Fault Protection Processors on classic systems provide opportunity for Ground Operations to fix problems with main flight processor. However, they are not powerful enough to run S/C by themselves, so not very useful for missions that must operate autonomously. • All processors not assigned to be the master act as “watchers”, hence more oversight than with a single FPP per flight processor • RF links between processors allows for N-fold redundancy • RF links between processors and subsystems allows for easier implementation of subsystem redundancy
- 34. A Realistic Interstellar Explorer Major Assumptions: X-band uplink and downlink Medium gain antenna on spacecraft (G= 15 dBic) 70m dish on ground, transmit power= 18.4 kW S/C transmit power= 0.5 W S/C receiver noise figure= 1.0 dB S/C passive loss= 1.0 dB Uplink 7.8 bps, 3 dB margin Uplink receiver lock threshold RSB- 5 Link Analysis Results -100.0 -125.0 dBm) UPLINK (Power -150.0 Received Downlink 10 bps, 3 dB margin -175.0 Total -200.0 -225.0 Earth Range (AU) Downlink receiver lock threshold DOWNLINK 1 10 100 1000
- 35. 00-0642-1 Optical Communication System System requirements: – Average transmit power > 20 W – Aperture: 1 meter – Burst data rate: 500 bps @ 1000 A.U. – Pointing accuracy ~ 300 nrad – Intensity modulation - direct detection – Co-boresighted fine guidance tracker – Off-axis coarse tracker JHU/APL concept incorporates advanced technologies (VCSELs, MEMs, and diffractive optics) to minimize mass and prime power Sparse Shack Hartmann array Fresnel objective Inver se Fresnel corrector Reimager Beamsplitter Off-axis coarse tracker light Focal plane array VCSEL array Beam steering feedback MEMs mirror Spatial deconvolution Outgoing laser light Beam former Beam shaping feedback Refocuser Tracker INS Incoming narrowband light S/C spin axis
- 36. Schedule 2000-2002 Advanced Technology Development study(ies) 2000-2002 Continued definition studies of the solar sail concept for IP at JPL 2002-2003 Update of OSS strategic plan with study for a "New Millennium"-like mission 2003-2007 Focused technology development for small probe technologies 2004-2007 Development of sail demonstration mission 2004-2007 Development of Solar Probe mission (test for perihelion propulsion) [2006-2007 Hardware tests for radioisotope sail feasibility ] [2006-2007 Hardware tests for antimatter propulsion schemes ] 2006-2007 Monitor DoD STP effort and conduct NASA-specific hardware tests [2002-2007 Development of space-qualified nuclear thermal reactor ] 2007-2010 Focused technology development for an Interstellar Probe 2009-2012 Design and launch of first generation solar-sail probe 2010 Test of Solar Probe performance in the perihelion pass of October 2010 2012-2015 Design and launch second generation probe 1000 AU goal in 50 years 2015-2065 Data return from 1000 AU and “beyond the infinite...”
- 37. Probes are Already En Route to Distant Stars Pioneer 10 as a relic, adrift and cold, passing through by a random star in the Milky Way (© Astronomy Magazine)
- 38. The Next Step ... What is still needed next is another factor of 10 in speed, to ... 200 AU yr-1, at which the first targeted interstellar crossing to Epsilon Eridani will take ~3400 years, the age of the Colossi of Memnon (Amehotep III - 18th dyn) Though not ideal, the stars would be within our reach
- 39. Implementing the Next Step The target terminal speed is 200 AU/yr = 948 km/s At an initial propellant fraction of 60%, the mass ratio is 2.5, the required specific impulse is 1.05x105s To maximize the specific impulse, the propellant of choice is again LH2 The specific impulse corresponds to an exhaust speed of 1035 km/s or H+ accelerated through ~5.6 kV x = gIsp m0 m˙ 1 - mfinal m0 ln m0 mfinal ÷ +1 æ è ç ö ø é ë ê ù ú û X is the distance traveled Issue is the sizing and specific mass of the power plant Some type of nuclear energy is required
- 40. Example System Assume 10 mg/s of H+ = 960 A of current => 5.35 MW of electrical power required Assume 50 year acceleration time => mpropellant = 15,800 kg m0 = 26,300 kg mfinal = 10,500 kg Assume 1.5 kg/kW => mpowerplant = 8000 kg => mpayload = 2500 kg During acceleration (50 years), probe travels 4250 AU Minimum size is set by reactor criticality, power processing, engines, propellant tansk Required power and H2 amounts comparable to manned Mars mission requirements
- 41. Ad Astra!