Advanced Lithium Ion Venus Explorer (ALIVE)

CD–2012–72 COMPASS Final Report: Advanced Lithium Ion Venus Explorer (ALIVE) Requestor: Michael Paul, Applied Research Laboratory / Pennsylvania State University COMPASS Lead: Steven R. Oleson COMPASS Concept Design Integration Lead: Melissa L. McGuire COMPASS Study System Integration Lead: Carl E. Sandifer II COMPASS Team Members Les Balkanyi Michael Bur Laura Burke Kristen Bury Tony Colozza John Dankanich Study Date: March 2012 Jon Drexler James Fincannon Jim Fittje John Gyekenyesi Geoffrey Landis Mike Martini Tom Packard Carlos Rodriguez Charles Sheehe Anita Tenteris Joe Warner Glenn Williams This work is preliminary in nature, capturing the progress of work in a design group. Advanced Lithium Ion Venus Explorer (ALIVE) Deleted this table before printing Version V2 V2 V2 V2 V3 V4 CD–2012–72 Date 5/31/2012 5/31/2012 5/31/2012 5/31/2012 6/15/2012 6/15/2012 Description Carl needs to update/format Science Carl needs to update/format Communications Carl needs to update Systems Carl needs to update formulas for Structures Planned completion for integration of technical details Out of Publishing, Lorie and Les made comments iii March 2012 COMPASS Final Report TABLE OF CONTENTS 1.0   2.0   3.0   4.0   5.0   INTRODUCTION ..........................................................................Error! Bookmark not defined.   STUDY BACKGROUND AND ASSUMPTIONS ......................................................................... 2   2.1   Introduction ......................................................................................................................... 2   2.1.1   Background/Past Potential Venus Missions (Needs to be updated and formattedces) ......................................................................................................................... 3   2.1.2   Report Perspective and Disclaimer ........................................................................ 4   2.2   Assumptions and Approach ................................................................................................ 4   2.3   Study Summary Requirements ........................................................................................... 6   2.3.1   Figures of Merit ..................................................................................................... 6   2.4   Growth, Contingency, and Margin Policy .......................................................................... 7   2.4.1   Terms and Definitions ........................................................................................... 7   2.4.2   Mass Growth .......................................................................................................... 9   2.4.3   Power Growth (Needs to be updated-ces) ........................................................... 10   2.5   Mission Description .......................................................................................................... 10   2.5.1   Mission Analysis Assumptions ............................................................................ 11   2.5.2   Mission Trades ..................................................................................................... 11   2.5.3   Mission ΔV Details .............................................................................................. 13   2.5.4   Mission Analysis Analytic Methods .................................................................... 13   2.5.5   Concept of Operations (CONOPS) ...................................................................... 13   2.5.6   Mission Communications Details ........................................................................ 16   2.6   Launch Vehicle Details ..................................................................................................... 18   2.6.1   Payload Fairing Configuration............................................................................. 19   BASELINE DESIGN ..................................................................................................................... 20   3.1   Top-Level Design ............................................................................................................. 20   3.1.1   Master Equipment List (MEL) ............................................................................ 20   3.1.2   S/C Total Mass Summary .................................................................................... 22   3.1.3   Power Equipment List (PEL) ............................................................................... 23   3.2   System-Level Summary (Needs to be updated-ces) ......................................................... 25   3.2.1   Propellant Calculations ........................................................................................ 25   AREAS FOR FUTURE STUDY ................................................................................................... 26   SUBSYSTEM BREAKDOWN ..................................................................................................... 27   5.1   Science Package ................................................................................................................ 27   5.1.1   Descent Instruments ............................................................................................. 27   5.1.2   Surface Instrument Details................................................................................... 28   5.1.3   Science Design and MEL .................................................................................... 29   5.2   Communications ............................................................................................................... 29   5.2.1   Communications Requirements ........................................................................... 29   5.2.2   Communications Assumptions ............................................................................ 30   5.2.3   Communications Design and MEL...................................................................... 33   5.2.4   Communications Recommendation ..................................................................... 34   5.3   Command and Data Handling ........................................................................................... 34   5.3.1   C&DH Requirements ........................................................................................... 34   5.3.2   C&DH Assumptions ............................................................................................ 34   5.3.3   C&DH Design and MEL ..................................................................................... 34   5.3.4   C&DH Trades ...................................................................................................... 36   5.3.5   C&DH Analytical Methods ................................................................................. 36   5.3.6   C&DH Risk Inputs............................................................................................... 36   5.3.7   C&DH Recommendation ..................................................................................... 36   5.4   Guidance, Navigation and Control ................................................................................... 37   iv Advanced Lithium Ion Venus Explorer (ALIVE) 5.4.1   GN&C Requirements ........................................................................................... 37   5.4.2   GN&C Assumptions ............................................................................................ 37   5.4.3   GN&C Design and MEL ..................................................................................... 37   5.4.4   GN&C Trades ...................................................................................................... 38   5.4.5   GN&C Analytical Methods ................................................................................. 38   5.4.6   GN&C Risk Inputs............................................................................................... 41   5.4.7   GN&C Recommendation ..................................................................................... 41   5.5   Electrical Power System ................................................................................................... 41   5.5.1   Power Requirements ............................................................................................ 41   5.5.2   Power Assumptions ............................................................................................. 42   5.5.3   Power Design and MEL ....................................................................................... 42   5.5.4   Power Trades ....................................................................................................... 46   5.5.5   Power Analytical Methods................................................................................... 46   5.5.6   Power Risk Inputs ................................................................................................ 46   5.5.7   Power Recommendation ...................................................................................... 46   5.6   Propulsion System ............................................................................................................ 46   5.6.1   Propulsion System Requirements ........................................................................ 46   5.6.2   Propulsion System Assumptions ......................................................................... 46   5.6.3   Propulsion System Design and MEL ................................................................... 47   5.6.4   Propulsion System Trades ................................................................................... 48   5.6.5   Propulsion System Analytical Methods ............................................................... 48   5.6.6   Propulsion System Risk Inputs ............................................................................ 49   5.6.7   Propulsion System Recommendation .................................................................. 49   5.7   Structures and Mechanisms .............................................................................................. 49   5.7.1   Structures and Mechanisms Requirements .......................................................... 49   5.7.2   Structures and Mechanisms Assumptions ........................................................... 49   5.7.3   Structures and Mechanisms Design and MEL ..................................................... 49   5.7.4   Structures and Mechanisms Trades ..................................................................... 52   5.7.5   Structures and Mechanisms Analytical Methods................................................. 52   5.7.6   Structures and Mechanisms Risk Inputs .............................................................. 53   5.7.7   Structures and Mechanisms Recommendation .................................................... 53   5.8   Thermal Control ................................................................................................................ 53   5.8.1   Cruise Deck Thermal Control .............................................................................. 54   5.8.2   Electric Heaters .................................................................................................... 54   5.8.1   MLI and Thermal Control Paint .......................................................................... 55   5.8.2   Radiator and Cold Plates...................................................................................... 56   5.9   Venus Atmospheric Environment ..................................................................................... 57   5.10   Aeroshell and Descent Thermal Control........................................................................... 58   5.10.1   Descent Electronics Enclosure Thermal Control ................................................. 59   5.11   Surface Lander Thermal Control ...................................................................................... 60   6.0   COST AND RISK .......................................................................................................................... 61   6.1   Cost ................................................................................................................................... 61   6.2   Risk ................................................................................................................................... 63   7.0   BIBLIOGRAPHY .......................................................................................................................... 68   APPENDIX A —ACRONYMS AND ABBREVIATONS ........................................................................ 70   APPENDIX B —RENDERED IMAGES................................................................................................... 72   B.1   Insert subtitle ..................................................................................................................... 72   APPENDIX C —COMPASS INTERNAL DETAILS (ALWAYS LAST) ............................................... 73   C.1   COMPASS Description .................................................................................................... 73   C.2   GLIDE Study Share .......................................................................................................... 73   C.2.1   GLIDE Architecture ............................................................................................ 73   CD–2012–72 v March 2012 COMPASS Final Report C.2.2   GLIDE Study Container ...................................................................................... 73   LIST OF FIGURES Figure 1.1—ALIVE S/C. .............................................................................................................................. 2   Figure 2.1—Previous Venus space vehicles. ................................................................................................ 3   Figure 2.2—Probe 5-d cruise and descent timeline ...................................................................................... 4   Figure 2.3—Cartesian Map of Ovda Regio. ................................................................................................. 5   Figure 2.4—Graphical illustration of the definition of basic, predicted, total and allowable mass. ............ 7   Figure 2.5—Trajectory graphic. Best case ALIVE opportunity. ................................................................ 11   Figure 2.6—Figure 2.6.2-1: Primary and backup mission opportunities.................................................... 11   Figure 2.7—Minimum arrival energy solution. .......................................................................................... 12   Figure 2.8—Example LGA option to reduce launch energy requirements. ............................................... 13   Figure 2.9—ALIVE EDL operations. ......................................................................................................... 15   Figure 2.10—Earth-Probe distance (a) and SEP and SPE angles (b). ........................................................ 16   Figure 2.11—SOAP communications analysis. .......................................................................................... 17   Figure 2.12—Need caption ......................................................................................................................... 17   Figure 2.13—Need caption ......................................................................................................................... 17   Figure 2.14— .............................................................................................................................................. 18   Figure 2.15— .............................................................................................................................................. 18   Figure 2.16— .............................................................................................................................................. 18   Figure 2.17—ALIVE Lander aeroshell dimensions. .................................................................................. 19   Figure 2.18—Isometric views of the ALIVE Lander inside the aeroshell. ................................................ 20   Figure 2.19—Landing legs and X-band antenna deployment sequence. .................................................... 20   Figure 3.1—ALIVE design approach. ........................................................................................................ 21   Figure 5.1—Block diagram of ALIVE communications hardware ............................................................ 31   Figure 5.2—Illustration of ALIVE LGA. ................................................................................................... 32   Figure 5.3—Graphic of ALIVE HGA. ....................................................................................................... 32   Figure 5.4—Image of ALIVE SDST communications hardware. .............................................................. 33   Figure 5.5—Image of ALIVE TWTA and EPC ......................................................................................... 33   Figure 5.6—Summary of nominal EDL profile. ......................................................................................... 39   Figure 5.7—Acceleration timeline from atmospheric entry. ...................................................................... 39   Figure 5.8—Nominal altitude profile during atmospheric entry. ............................................................... 40   Figure 5.9 Duplex Sketch............................................................................................................................ 43   Figure 5.11—Heat and power flows for ALIVE Power and Cooling System. ........................................... 44   Figure 5.12—ALIVE Power/Cooling System highlights. .......................................................................... 45   Figure 5.13—Feed System components. .................................................................................................... 47   Figure 5.14—Preliminary Cruse Deck Propulsion P&ID ........................................................................... 47   Figure 5.15—(a) The Lander stowed within the heat shield/backshell assembly and (b) the Lander fully deployed. ........................................................................................................................................ 50   Figure 5.16—DuPont Kapton Strip Heater. ................................................................................................ 54   Figure 5.17—Example of MLI blanket design and application.................................................................. 55   Figure 5.18—Example of a cold plate with integrated heat pipes. ............................................................. 56   Figure 5.19—Radiator with integral heat pipes (ACT, inc). ...................................................................... 57   Figure 5.20—Venus atmospheric properties. ............................................................................................. 57   Figure 5.21—Venus atmospheric structure. ............................................................................................... 58   Figure 5.22—Orion heat shield structural makeup. .................................................................................... 59   Figure 5.23—Stardust Aeroshell Geometry................................................................................................ 59   Figure 5.24—Descent electronics thermal control items. ........................................................................... 60   Figure 5.25—Heat leak into the Lander electronics enclosure pressure vessel. ......................................... 61   vi Advanced Lithium Ion Venus Explorer (ALIVE) Figure 6.1—ALIVE Risk List. .................................................................................................................... 64   Figure 6.2—Risks 1 and 2—Mission and Power risks for ALIVE. ........................................................... 65   Figure 6.2—Risks 3 and 4—Mechanisms and Mission risks for ALIVE. ................................................. 65   Figure 6.3—Risks 5 and 6—Thermal and Mission risks for ALIVE. ........................................................ 66   Figure 6.4—Risks 7 and 8—Electronics and Structures risks for ALIVE. ................................................ 66   Figure 6.5—Risks 9 and 10—Thermal risks for ALIVE. ........................................................................... 67   Figure 6.6—Risks 11 and 12—Thermal and Propulsion risks for ALIVE. ................................................ 67   Figure 6.7—ALIVE TRL assessment. ........................................................................................................ 68   Figure B.1— ................................................................................................................................................ 72   Figure B.2— ................................................................................................................................................ 72   Figure B.3— ................................................................................................................................................ 72   LIST OF TABLES Table 1.1—Mission and S/C Summary for the ALIVE mission .................................................................. 2   Table 2.1—Assumptions and Study Requirements ...................................................................................... 5   Table 2.2—Definition of Masses Tracked in the MEL ................................................................................ 9   Table 2.3—MGA and Depletion Schedule (AIAA S-120-2006) ............................................................... 10   Table 2.4—LV performance versus launch energy of interest. .................................................................. 12   Table 2.5—Mission ΔV Summary for the ALIVE S/C .............................................................................. 13   Table 2.6—Additional Mission Analysis.................................................................................................... 13   Table 3.1—ALIVE MEL WBS Format ...................................................................................................... 21   Table 3.2—ALIVE System Summary ........................................................................................................ 22   Table 3.3—ALIVE Total Mass With Payload (Includes 30% System Level Growth) .............................. 23   Table 3.4—Definition of the ALIVE S/C Power Modes ............................................................................ 23   Table 3.5—ALIVE S/C PEL ...................................................................................................................... 24   Table 3.6—Case x Thermal Waste Heat Per Power Mode ......................................................................... 24   Table 3.7—ALIVE S/C Propellant Details ................................................................................................. 25   Table 3.8—Inert Mass Calculations For ALIVE Total S/C ........................................................................ 26   Table 3.9—ALIVE Architecture Details .................................................................................................... 26   Table 5.1—Descent Instruments ................................................................................................................. 27   Table 5.2—Surface Instruments ................................................................................................................. 28   Table 5.3—Science ALIVE MEL ............................................................................................................... 29   Table 5.4—Communications Science Link Budget .................................................................................... 30   Table 5.5—Communications Case x MEL ................................................................................................. 33   Table 5.6—C&DH ALIVE S/C MEL......................................................................................................... 35   Table 5.7—GN&C ALIVE S/C MEL......................................................................................................... 37   Table 5.8—Assumptions Made During Parachute Sizing .......................................................................... 40   Table 5.9—Power Requirements ................................................................................................................ 42   Table 5.10 Mass Breakdown of Duplex Power System .............................................................................. 45   Table 5.11—Electrical Power System ALIVE S/C MEL ........................................................................... 45   Table 5.12—Propulsion System ALIVE S/C MEL .................................................................................... 48   Table 5.13—ALIVE S/C Structures MEL .................................................................................................. 50   Table 5.14—Lander Structures MEL .......................................................................................................... 51   Table 5.15—Aeroshell Structures MEL ..................................................................................................... 51   Table 5.16—Cruise Deck Structures MEL ................................................................................................. 52   Table 5.17—Transit Environment Constants.............................................................................................. 54   Table 5.18—Thermal ALIVE S/C MEL ..................................................................................................... 55   Table 5.19—MLI Specifications................................................................................................................. 56   Table 5.20—Cruise Deck Radiator Sizing .................................................................................................. 57   Table 5.21—Heatshield Material Layer Properties .................................................................................... 59   Table 5.22—Insulation and Phase Change Material Specifications ........................................................... 60   CD–2012–72 vii March 2012 COMPASS Final Report Table 5.23—Pressure Vessel Components and Heat Leak ......................................................................... 61   Table 6.1—COMPASS Subsystem Level Cost Breakdown—ALIVE ....................................................... 62   Table 6.2—Lifecycle Cost Comparison for the ALIVE mission ................................................................ 63   Table C.1—Study Container and Descriptions ........................................................................................... 73   viii Advanced Lithium Ion Venus Explorer (ALIVE) STUDY PARTICIPANTS (NEEDS TO BE UPDATED-CES) Study Name Design Session Subsystem Name Center Michael Paul PSU James Kasting PSU Propulsion PI George Schmidt GRC Science PI, Robotic Elements Geoffrey Landis GRC Gary Hunter GRC Propulsion PI Science PI Robotic Elements Email COMPASS Team Lead Steve Oleson GRC [email protected] System Integration, MEL, Mission Visualization, and Final Report Documentation Carl E. Sandifer II GRC [email protected] Technical Editing and Oversight Melissa.L.Mcguire GRC [email protected] Internal Editing and Final Report Documentation Leslie Balkanyi GRC [email protected] Mission John Dankanich, GRC [email protected] Mission Ian Dux GRC [email protected] Michael Martini QNA Corp [email protected] ELV, Integration and Test, Operations Carlos Rodriguez GRC [email protected] Propulsion and Propellant James Fittje QinetiQ [email protected] Propulsion and Propellant David Chato GRC [email protected] Mechanical Systems John Gyekenyesi GRC [email protected] Mechanical Systems David McCurdy ASRC [email protected] Thermal Tony Colozza QNA Corp [email protected] Power Paul Schmitz PCS [email protected] Power Timothy Miller PSU Glenn L. Williams GRC [email protected] Charles Sheehe GRC [email protected] Tom Packard GRC [email protected] Jonathan Drexler GRC Anita Tenteris GRC Mission, Operations, GN&C Command and Data Handling Communications Configuration Cost Risk/Reliability CD–2012–72 ix [email protected] March 2012 COMPASS Final Report x Advanced Lithium Ion Venus Explorer (ALIVE) 1.0 EXECUTIVE SUMMARY The COncurrent Multidisciplinary Preliminary Assessment of Space Systems (COMPASS) Team partnered with the Applied Research Laboratory to perform a NASA Innovative Advanced Concepts (NIAC) Program study to evaluate chemical based power systems for keeping a Venus lander alive (power and cooling) and functional for a period of days. The mission class targeted was either a Discovery ($500M) or New Frontiers ($750M to $780M) class mission. Historic Soviet Venus landers have only lasted on the order of 2 hours in the extreme Venus environment: temperatures of 460 °C and pressures of 93 bar. Longer duration missions have been studied using plutonium powered systems to operate and cool landers for up to a year. However, the plutonium load is very large. This NIAC study sought to still provide power and cooling but without the plutonium. Battieries are far too heavy but a system which uses the atmosphere (primarily carbon dioxide) and on onboard fuel to power a power generation and cooling system was sought. The resuling design was the Advanced Long-Life Lander Investigating the Venus Environment (ALIVE) Spacecraft (S/C) which burns lithium (Li) with the CO2 atmosphere to heat a Duplex Stirling to power and cool the lander for a 5 day duration (until the Li is exhausted). While it does not last years a chemical powered system surviving days eliminates the cost associated with utilizing a flyby relay S/C and allows a continuous low data rate direct to earth (DTE) link in this instance from the Ovda Regio of Venus. The five day collection time provided by the chemical power systems also enables science personnel on earth to interact and retarget science – something not possible with a ~2 hour spacecraft lifetime. It also allows for contingency operations directed by the ground (reduced risk). The science package was based on that envisioned by the Venus Intrepid Tessera Lander (VITaL) Decadal Survey Study. The Li Burner within the long duration power system creates approximately 14000 W of heat. This 1300 °C heat using Li in the bottom ‘ballast’ tank is melted to liquid by the Venus temperature, drawn into a furnace by a wick and burned with atmospheric CO2.. The Li carbonate exhaust is liquid at 1300°C and being denser than Li drains into the the Li tank and solidifies. Since the exhaust product is a dense liquid no ‘chimney’ is required which conserves the heat for the stirling power convertor. The Duplex Stirling provides about 300 W of power and removes about 300 W of heat from the avionics and heat that leaks into the 1 bar insulated payload pressure vessel kept at 25 °C. The NaK radiator is run to the top of the drag flap. The ALIVE vehicle is carried to Venus via an Atlas 411 launch vehicle (LV) with a C3 of 7 km2/s2. An Aeroshell, derived from the Genesis mission, enables a direct entry into the atmosphere of Venus (–10°, 40 g max) and 6 m/s for landing (44 g) using a drag ring. For surface science and communication, a 100 WRF, X-Band 0.6m pointable DTE antenna provides 2 kbps to DSN 34 m antenna clusters. Table 1.1 summarizes the top-level details of each subsystem that was incorporated into the design. Cost estimates of the ALIVE mission show it at ~ $760M which puts it into the New Frontiers class. The ALIVE landed duration is only limited by the amount of Li which can be carried by the lander. Further studies are needed to investigate how additional mass can be carried, perhaps by a larger launcher and larger aeroshell. Other power conversion/cooling systems might also bring other benefits. CD–2012-72 1 March 2012 COMPASS Final Report Figure 1.1—ALIVE S/C. Table 1.1—Mission and S/C Summary for the ALIVE mission Subsystem area Top-level system Mission and operations, and Guidance, Navigation and Control (GN&C) Launch Science Power Propulsion Structures and mechanisms Communications Command and Data Handling (C&DH) Thermal Details 5-d Venus lander for scientific explorer of Venus, Mass Growth per to AIAA S-1202006 (add growth to make system level 30%) Direct to Venus, genesis aeroshell, parachute to remove aeroshell and backshell Atlas 411 class Landed and descent science packages similar to VITaL 2010 Decadal survey study. Landed science Pan Cam, context imager, and LIBS for in-situ science Li/Atm CO2 burner, Duplex Stirling power (300 We)/Cooling (300 W-hr), Li tank also used as ballast, sodium-potassium alloy (NaK) radiator placed on drag flag, high temperature sodium-sulfur (NaS) batteries for load leveling Hydrazine monopropellant for RCS and mid course corrections ~ 5g launch, 40 g entry and landing loads, all metallic, pressure vessels to handle 90 bar Venus atmospheric pressure Waveguide with window between the coldbay and external antenna. Omni antennas for telemetry/control during cruise/descent 2 kbps data rates for landed science, 1 GB storage, 100 WRF X-band DTE 0.6 m pointable antenna. External Venus temps 90 bar/460 °C max, Internal vault pressure/temps 1 bar/ 25 °C max 2.0 STUDY BACKGROUND AND ASSUMPTIONS 2.1 Introduction Total Lander mass with growth (kg) 167 47 311 606 53 30 42 NIAC has sponsored an effort to evaluate chemical based power systems by keeping a Venus lander alive (power and cooling) for a period of 5 days. The ALIVE S/C consists of three elements: the Cruise Deck, Aeroshell, and Lander. The Cruise Deck is responsible for housing the hydrazine monopropellant for the reaction control system (RCS) and for mid-course corrections after separating from the Atlas 411 Expendable Launch Vehicle (ELV). The Aeroshell enables a direct entry into the atmosphere of Venus (–10°, 40 g max). The 2 Advanced Lithium Ion Venus Explorer (ALIVE) Aeroshell is jettisoned after the Lander parachute is deployed to allow for a secure landing with the support of a fixed drag flap to reduce the landing velocity. The Lander is designed to operate within a 460 °C (860 °F) environment with a pressure of 93 bar (9,300,000 Pa) while sized to support surface science and communications with the Earth-based DSN for 5 days. Assuming the targeted landing site of Ovda Regio, the science objectives include: § § § § Correlating high altitude mountain surface reflectivity from radar measurements with surface data Investigating mineralogy and weathering of the Venus surface Evaluating the past extent of Venus oceans Increasing knowledge of Venus weather From a cost perspective, the drive was to design a S/C that will meet the requirements of a Discovery ($500M) or New Frontiers ($780M) class mission. 2.1.1 Background/Past Potential Venus Missions (Needs to be updated and formatted-ces) Referenced from the VITaL mission concept study report, the Russian Venera Landers utilized lithium nitrate trihydrate (LNT) for phase change material to provide maximum conduction to electronics. There were 10 Venera probes that successfully landed on the surface of Venus and transmitted data between 1964 and 1982 (Balint, Tibor). The U.S. Pioneer Venus mission of 1978 operated similarly to the Venera Landers. Typically, these lenders survived for less than an hour on the surface due to the harsh environment. Figure 2.1 shows a variety of probes previously sent to explore Venus. The VITaL mission from the recent Decadal survey is comparable to the ALIVE science objectives. Figure 2.2 shows a typical entry, descent, and landing (EDL) timeline for a Venus lander. Figure 2.1—Previous Venus space vehicles. CD–2012-72 3 March 2012 COMPASS Final Report Figure 2.2—Probe 5-d cruise and descent timeline 2.1.2 Report Perspective and Disclaimer This report is meant to capture the study performed by the COMPASS Team, recognizing that the level of effort and detail found in this report will reflect the limited depth of analysis that was possible to achieve during a concept design session. All of the data generated during the design study is captured within this report in order to retain it as a reference for future work. 2.2 Assumptions and Approach The harsh environment of Venus provides a number of challenges in the operation of equipment and materials. Operating within this environment, from entry to descent to operation on the surface requires significant thermal control. The atmosphere is composed of mainly CO2 but does contain corrosive components such as sulfuric acid. The planet has a very thick atmosphere and is completely covered with clouds. The temperature and pressure near the surface is 455 °C at 90 bar. The Ovda Regio location on Venus was chosen to be the landing and surface science location to maximize communication with Earth, while providing a high altitude for science reflectivity. A Cartesian map of Ovda Regio can be found in Figure 2.3. The assumptions and requirements about the ALIVE S/C, including those that were known prior to starting the COMPASS design study session, are shown in Table 2.1. This table gathers the assumptions and requirements and calls out trades that were considered during the course of the design study, and offthe-shelf (OTS) materials that were used wherever possible. 4 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 2.3—Cartesian Map of Ovda Regio. Table 2.1—Assumptions and Study Requirements Subsystem area Top-level Requirements/Assumptions 5 day Venus lander for scientific exploration of Venus. Science investigations based on VITaL FOMs: Surface duration, Science collected, Science data returned, Cost Trades Science approach, duration, fuel System Identify new technologies, TRL 6 cutoff 2018, 2022 launch year, single fault tolerant. Earth directed operations for 5 d on Venus surface Mass growth per to AIAA S-120-2006 (add growth to make system level 30%) Mission and operations and GN&C Direct to Venus, C3 = 7 km /s , –10° entry angle, genesis aeroshell, parachute to remove aeroshell and backshell, fixed drag-flap to slow landing speed to 6 m/s LV Atlas 411 class Launch Loads: Axial SS ± 4.5 g, Lateral ± 1g Science Landed and descent science packages similar to VITaL 2010 Decadal Placement of instruments, number of survey study. Descent science in separate pressure vessel to images minimize landed pressure vessel (atm spectrometers and imagers). Landed science Pan Cam, context imager, and LIBS for in-situ science Propulsion Hydrazine monopropellant for RCS and mid course corrections Biprop for starting at GTO Power Li/Atm CO2 burner, Duplex Stirling power (300 We) / Cooling (300 Whr), Li tank also used as ballast, NaK radiator placed on drag flag, high temperature NaS batteries for load leveling Brayton or Stirling, Fuel type (Li, MgAl), batteries (NaS), power convertor/cooler C&DH/ Communications 2 kbps data rates for landed science, 1 GB storage, 100 WRF X-band DTE 0.6 m pointable antenna. Waveguide with window between the coldbay and external antenna. Omni antennas for telemetry/control during cruise/descent Bluetooth controllers to eliminate feedthroughs, Data storage, MIPS, operating temperature, Pointing, data rate (2 kbps), store/deploy CD–2012-72 2 2 5 Parachute descent time, aeroshell sizing, deceleration gs, ballutes. Aeroshell (3.6 m, Genesis derivative) and parachute (descent time) sizing. Parachute separation March 2012 COMPASS Final Report Subsystem area Thermal and environment Requirements/Assumptions Trades External Venus temps 90 bar/460 °C max, Internal vault pressure/temps 1 bar/25 °C max. 20 cm aerogel insulation inside internal vault, avionics waste head and heat leak (~300 Wth) removed with Stirling cooler. 3.6 m Aeroshell base on -10° entry angle and Genesis Internal pressure (ambient vs 1 atm vs vacuum) and insulation (aerogel or MLI), windows for science and comms, minimize wire feedthroughs, active sterling or passive pre-use of chemical fuel to absorb surface heat in 25°C temperature, aeroshell Mechanisms Deployable Legs with crushable pads, deployable, pointable X-band antenna, Aeroshell and cruise deck separations Number, size of wheels Structures ~ 5g launch, 40 g entry and landing loads, all metallic, pressure vessels to handle 90 bar Venus atmospheric pressure What pressure for cold box? Trade 1 bar vs 90 bar S/C, reuse pressure vessel as aeroshell Cost New Frontiers Assumptions, 2015 $ Discovery and New Frontiers assumptions Risk Major Risks: high temp mechanisms/gimbals, landing 2.3 Study Summary Requirements 2.3.1 Figures of Merit (Provided by System Integration Lead) The following are the figures of merit (FOM) and/or the elements upon which the design is judged to assess the closure of the study and whether or not the design meets the requirements of the customer: § § § § § § Mass: Must fit inside an Atlas V 411 Launch Date: 2023 (primary), 2024 (backup) Reliability: Single fault tolerant (where applicable) Cost: Discovery or New Frontiers class Effectiveness/applicability/flexibility of chemical power system Lifetime and survivability on Venus Surface 6 Advanced Lithium Ion Venus Explorer (ALIVE) 2.4 Growth, Contingency, and Margin Policy (Basic = bottoms-up estimate of dry mass. Mass growth allowances (MGA) = applied per subsystem line item) Figure 2.4—Graphical illustration of the definition of basic, predicted, total and allowable mass. 2.4.1 Terms and Definitions Mass Basic Mass (aka CBE Mass) CBE Mass Dry Mass Wet Mass CD–2012-72 The measure of the quantity of matter in a body. Mass data based on the most recent baseline design. This is the bottoms-up estimate of component mass, as determined by the subsystem leads. Note 1: This design assessment includes the estimated, calculated, or measured (actual) mass, and includes an estimate for undefined design details like cables, MLI, and adhesives. Note 2: The MGA and uncertainties are not included in the basic mass. Note 3: COMPASS has referred to this as current best estimate (CBE) in past mission designs. Note 4: During the course of the design study, the COMPASS Team carries the propellant as line items in the propulsion system in the Master Equipment List (MEL). Therefore, propellant is carried in the basic mass listing, but MGA is not applied to the propellant. Margins on propellant are handled differently than they are on dry masses. See Basic Mass. The dry mass is the total mass of the system or S/C when no propellant is added. The wet mass is the total mass of the system, including the dry mass and all of the propellant (used, predicted boil-off, residuals, reserves, etc.). It should be noted that in human S/C designs the wet masses would include more than propellant. In these cases, instead of 7 March 2012 COMPASS Final Report Inert Mass Basic Dry Mass CBE Dry Mass MGA Predicted Mass Predicted Dry Mass Mass Margin (aka Margin) System-Level Growth propellant, the design uses Consumables and will include the liquids necessary for human life support. In simplest terms, the inert mass is what the trajectory analyst plugs into the rocket equation in order to size the amount of propellant necessary to perform the mission delta-Velocities (ΔVs). Inert mass is the sum of the dry mass, along with any non-used, and therefore trapped, wet materials, such as residuals. When the propellant being modeled has a time variation along the trajectory, such as is the case with a boil-off rate, the inert mass can be a variable function with respect to time. This is basic mass (aka CBE mass) minus the propellant or wet portion of the mass. Mass data is based on the most recent baseline design. This is the bottoms-up estimate of component mass, as determined by the subsystem leads. This does not include the wet mass (e.g., propellant, pressurant, cryo-fluids boil-off, etc.). See Basic Dry Mass. MGA is defined as the predicted change to the basic mass of an item based on an assessment of its design maturity, fabrication status, and any in-scope design changes that may still occur. This is the basic mass plus the mass growth allowance for to each line item, as defined by the subsystem engineers. Note : When creating the MEL, the COMPASS Team uses Predicted Mass as a column header, and includes the propellant mass as a line item of this section. Again, propellant is carried in the basic mass listing, but MGA is not applied to the propellant. Margins on propellant are handled differently than they are handled on dry masses. Therefore, the predicted mass as listed in the MEL is a wet mass, with no growth applied on the propellant line items. This is the predicted mass minus the propellant or wet portion of the mass. The predicted mass is the basic dry mass plus the mass growth allowance as the subsystem engineers apply it to each line item. This does not include the wet mass (e.g., propellant, pressurant, cryo-fluids boil-off, etc.). This is the difference between the allowable mass for the space system and its total mass. COMPASS does not set a Mass Margin, it is arrived at by subtracting the Total mass of the design from the design requirement established at the start of the design study such as Allowable Mass. The goal is to have Margin greater than or equal to zero in order to arrive at a feasible design case. A negative mass margin would indicate that the design has not yet been closed and cannot be considered feasible. More work would need to be completed. The extra allowance carried at the system level needed to reach the 30% aggregate MGA applied growth requirement. For the COMPASS design process, an additional growth is carried and applied at the system level in order to maintain a total growth on the dry mass of 30%. This is an internally agreed upon requirement. Note 1: For the COMPASS process, the total growth percentage on the basic dry mass (i.e. not wet) is: 8 Advanced Lithium Ion Venus Explorer (ALIVE) Total Growth = System Level Growth + MGA*Basic Dry Mass Total Growth = 30%* Basic Dry Mass Total Mass = 30%*Basic Dry Mass + basic dry mass + propellants. Note 2: For the COMPASS process, the system level growth is the difference between the goal of 30% and the aggregate of the MGA applied to the Basic Dry Mass. MGA Aggregate % = (Total MGA mass/Total Basic Dry Mass)*100 Where Total MGA Mass = Sum of (MGA%*Basic Mass) of the individual components System Level Growth = 30%* Basic Dry Mass – MGA*Basic Dry Mass = (30% – MGA aggregate %)*Basic Dry Mass Note 3: Since CBE is the same as Basic mass for the COMPASS process, the total percentage on the CBE dry mass is: Dry Mass total growth +dry basic mass = 30%*CBE dry mass + CBE dry mass. Total Mass Allowable Mass Therefore, dry mass growth is carried as a percentage of dry mass rather than as a requirement for LV performance, etc. These studies are Pre-Phase A and considered conceptual, so 30% is standard COMPASS operating procedure, unless the customer has other requirements for this total growth on the system. The summation of basic mass, applied MGA, and the system-level growth. The limits against which margins are calculated. Note: Derived from or given as a requirement early in the design, the allowable mass is intended to remain constant for its duration. Table 2.2 expands definitions for the MEL column titles to provide information on the way masses are tracked through the MEL used in the COMPASS design sessions. These definitions are consistent with those above in Figure 2.4 and in the terms and definitions. This table is an alternate way to present the same information to provide more clarity. Table 2.2—Definition of Masses Tracked in the MEL CBE mass Mass data based on the most recent baseline design (includes propellant) CBE dry + propellant 2.4.2 MGA growth Predicted change to the basic mass of an item phrased as a percentage of CBE dry mass MGA% * CBE dry = growth Predicted mass Predicted dry mass The CBE mass plus the mass The CBE mass plus the mass growth allowance (MGA) — growth allowance (MGA) propellant CBE dry + propellant + growth CBE dry + growth Mass Growth The COMPASS Team normally uses the AIAA S–120–2006, “Standard Mass Properties Control for Space Systems,” as the guideline for its mass growth calculations. Table 2.3 shows the percent mass growth of a piece of equipment according to a matrix that is specified down the left-hand column by level of design maturity and across the top by subsystem being assessed. The COMPASS Team’s standard approach is to accommodate for a total growth of 30% or less on the dry mass of the entire system. The percent growth factors shown above are applied to each subsystem before an additional growth is carried at the system level, in order to ensure an overall growth of 30%. Note that for designs requiring propellant, growth in the propellant mass is either carried in the propellant calculation itself or in the ΔV used to calculate the propellant required to fly a mission. CD–2012-72 9 March 2012 COMPASS Final Report The system-integration engineer carries a system-level MGA, called “margin”, in order to reach a total system MGA of 30%. This is shown as the mass growth for the allowable mass on the authority to precede line in mission time. After setting the margin of 30% in the preliminary design, the rest of the steps shown below are outside the scope of the COMPASS Team. Table 2.3—MGA and Depletion Schedule (AIAA S-120-2006) A 6 7 2.4.3 Wire harness Instrumentation ECLSS, crew systems 5 Propulsion 4 Mechanisms C Thermal control 3 >15 kg Solar array 2 5 to 15 kg Battery E 0 to 5 kg Brackets, clips, hardware 1 Electrical/electronic components Structure Maturity code Major category MGA (%) 30 25 20 25 30 25 30 25 25 25 55 55 23 25 20 15 15 20 15 20 20 15 15 30 30 15 20 15 10 10 15 10 10 15 10 10 25 25 10 10 5 5 5 6 5 5 5 5 5 10 10 6 3 3 3 3 3 3 3 2 3 3 5 5 4 Design maturity (basis for mass determination) Estimated (1) An approximation based on rough sketches, parametric analysis, or undefined requirements; (2) A guess based on experience; (3) A value with unknown basis or pedigree Layout (1) A calculation or approximation based on conceptual designs (equivalent to layout drawings); (2) Major modifications to existing hardware Prerelease designs (1) Calculations based on a new design after initial sizing but prior to final structural or thermal analysis; (2) Minor modification of existing hardware Released designs (1) Calculations based on a design after final signoff and release for procurement or production; (2) Very minor modification of existing hardware; (3) Catalog value Existing hardware (1) Actual mass from another program, assuming that hardware will satisfy the requirements of the current program with no changes; (2) Values based on measured masses of qualification hardware Actual mass Measured hardware Customer furnished equipment or specification value No mass growth allowance—Use appropriate measurement uncertainty values Typically a “not-to-exceed” value is provided; however, contractor has the option to include MGA if justified Power Growth (Needs to be updated-ces) The COMPASS Team typically uses a 30% margin on the bottoms-up power requirements of the bus subsystems when modeling the amount of required power. Table 3.5 (Section 3.1.3) shows the power system assumptions specific to this design study. 2.5 Mission Description The baseline mission is a launch on May 18, 2023, direct from Earth to Venus. The mission does not require any deterministic post launch ΔV and only requires a launch energy of 6.2 km2/s2. The interplanetary transit is 160 d and arrives on October 24, 2023, with an arrival V∞ of approximately 4 km/s. 10 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 2.5—Trajectory graphic. Best case ALIVE opportunity. 2.5.1 Mission Analysis Assumptions The data provided from mission is prior to performance margin consideration. An additional 10% of LV performance will be decremented at the system level. Because there are no deep space maneuvers, no additional margin is included. 2.5.2 Mission Trades The mission evaluation included a performance assessment over potential launch opportunities from 2020 to 2025. Because the transfer to Venus does not require deterministic post launch ΔV, the launch energy is the only driver in Venus arrival mass capability. Over the launch window, the higher performance launch opportunity and backup dates are May 18, 2023, and December 25, 2024. The S/C and LV capability must be constrained to accommodate either opportunity. May 18, 2023, is the first and therefore baseline mission, however; the LV capability must accommodate the slight energy increase for the backup. The primary and backup missions are illustrated in Figure 2.6. Figure 2.6—Primary and backup mission opportunities. The examples in Figure 2.6 are for a Falcon 9 Block 2, however; the required launch energy is independent of the LV. The goal was to fit the S/C onto a Falcon 9. Unfortunately the final arrival mass requirements moved the mission onto an EELV class vehicle. The performance of the LV options considered is shown in Table 2.4. The length of the launch window was also evaluated. A 2-week launch window can be accommodated with a launch energy margin of only 0.1 km2/s2 and a 3-wk launch window CD–2012-72 11 March 2012 COMPASS Final Report can be accommodated with launch energy margin of 0.5 km2/s2; 6.65 km2/s2 is required for the baseline launch energy with a 3-wk launch window. Table 2.4—LV performance versus launch energy of interest. 2 C3, km /s 5 7 9 11 13 15 2 Falcon 9 2145 2015 1890 1765 1650 1540 Launch mass (kg) Atlas V 401 2720 2600 2480 2365 2255 2145 Atlas V 411 3550 3400 3255 3115 2980 2845 The launch energy for the primary and backup missions is 6 to 7 km2/s2, however; an option to launch with higher launch energy to minimize the arrival energy was also explored. The baseline mission has an arrival energy of 15.4 km2/s2, the highest of any mission option. There is a small range where the arrival launch energy can be reduced while still requiring no deep space maneuvers. Minimizing the arrival energy will change the launch opportunity slightly. Because the mission did not close on a Falcon 9, there is significant margin and virtually no penalty in launching to the higher C3 and reducing the entry system requirements. An example solution minimizing the arrival energy is shown in Figure 2.7. Figure 2.7—Minimum arrival energy solution. Another option evaluated by not selected was to use a lunar gravity assist (LGA) in order to attempt to stay on the Falcon 9 (Figure 2.8). The only viable option to reduce the LV requirement for a trajectory to Venus is to launch to a negative C3 and leverage an LGA. Using a launch energy less than escape and performing maneuvers for the LGA and powered deep gravity well burn at Earth, the delivered mass capability of the Falcon 9 can be increased. The LGA does increase the Falcon 9 capability from ~2,000 kg to over 2,500 kg to Venus, it does require a large propulsion system. It was preferred to baseline a larger and higher cost LV rather than accept the increased S/C complexity and cost. 12 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 2.8—Example LGA option to reduce launch energy requirements. (Note: This is a Mars example.) 2.5.3 Mission ΔV Details Table 2.4 shows a ΔV summary throughout the mission. The vast majority of the ΔV is used for trajectory correction maneuvers (TCM). Analysis of the amount of ΔV used by the MESSENGER S/C revealed that less than 40 m/s of ΔV was used before the S/C first flew by Venus on its way to Mercury, hence it was assumed that ALIVE would need roughly 40 m/s of ΔV for TCMs on its way to Venus. An Isp of 220 s was assumed for the propulsion system. Table 2.5—Mission ΔV Summary for the ALIVE S/C Phase no. 1 2 3 4 5 Phase name Null tip-off rates TCM 1 TCM 2 Spin-up Separation Pre-burn mass (kg) 2478 2477 2454 2431 232 ΔV (m/s) 1 20 20 2 2 Prop used (kg) 1.1 22.9 22.6 2.3 0.2 Post burn mass (kg) 2477 2454 2431 2429 232 Table 2.6—Additional Mission Analysis Insert table here 2.5.4 Mission Analysis Analytic Methods For the mission design of the ALIVE mission, both Mission Analysis Low-Thrust Optimization (MALTO) and Copernicus were used for trajectory design. The MALTO program was used in ΔV mode for ballistic trajectory optimization. MALTO can only be used for the interplanetary mission design. Copernicus was also used for minimum ΔV optimization of the interplanetary transfer and landing site targeting. 2.5.5 Concept of Operations (CONOPS) (1) Pre-Launch Ops and Cruise to Venus ALIVE will be launched from the NASA Kennedy Space Center (KSC) on an Atlas V 411, which will carry all the elements necessary for the mission. The launch date for the analysis is May 18, 2023. Payload will be switched to internal power 5 min before liftoff and will remain on battery power until solar array deployment at approximately 1.5 hr MET. CD–2012-72 13 March 2012 COMPASS Final Report There is the potential for launch safety concerns due to the presence of solid Li, which is needed for the payload’s Stirling engine operation. These concerns will need to be identified and addressed separately, but given the experience of the U.S. Navy in successfully handling solid Li/Rankine torpedo systems we do not foresee any insurmountable difficulties. The Atlas upper stage will put ALIVE on a trans-Venus injection trajectory roughly 1.5 hr after liftoff. The solar arrays will then be deployed, allowing the S/C to generate its own power. ALIVE will immediately go through a complete vehicle assessment and the first of several instrument testing and calibration sessions. Communications with DSN during the cruise portion of the mission will be through the X-Band omni directional antennas located on the S/C aeroshell. Two hydrazine tanks and 16 thrusters will provide RCS propulsion and control. The cruise to Venus will last 159.6 d. (2) Arrival, Entry, Separation, and Lander Descent At Entry-20 min (E –20 min) the ALIVE S/C will be maneuvered to entry-attitude and the Lander’s beacon turned on. Shortly after, the descent instruments will be activated for science mode. At E –15 min the vehicle will be spun-up to 12 rpm, 5 min later the Lander will separate from the cruise deck, which will subsequently begin a divert burn collision avoidance maneuver (CAM). Communications with DSN will still be performed through the aeroshell X-band omni antennas. The Lander will go beacon-only as it enters the Venus atmosphere at an angle of –8.7° and an altitude of ~200 km. At about 90 km altitude, or 1.6 min after entry, ALIVE begins its descent science operations. At 65 km the subsonic parachute is deployed and the heat shield is released. Immediately after, the landing legs of the Lander are deployed. The parachute is released 20 min after deployment and the aeroshell departs with it. Communication with DSN is now through the X-Band omni directional antennas located on the Lander. After approximately 70 min of free-fall, ALIVE will land on the Venus surface, at less than 10 m/s and ~40 g’s. (3) Descent Science After entering the Venus atmosphere, and starting at about 90 km altitude, the Lander descent science instruments begin operating and storing data. This portion of the mission will last about 1.5 hr. The descent data is scheduled for transmission back to Earth during landed operations. For this analysis we assumed four principal science instruments used during descent: § Atmospheric Structure Investigation (ASI).—Starting at 90 km altitude, the ASI will make ten 12-b measurements every 10 m, for a total of 1.1 Mb of data, compressed at 10:1 § Neutral Mass Spectrometer (NMS).—The NMS begins gathering data at 30 km and will do 300 measurements before landing, capturing 1.8 Mb of data § Tunable Laser Spectrometer (TLS).—Beginning also at 30 km the TLS will also do 300 measurements during descent, or 3.6 Mb of data. § Descent Imager – Used only during the last 10 km of descent, it will capture 20 images for a total of 96 Mb of data (LOCO compressed) The expected total science data volume gathered by these instruments during descent should be approximately 105 Mb. (4) Early Landed Operations ALIVE is designed to operate for 5 consecutive days (or 120 hr) after landing on the surface of Venus. The first major operation is the ignition of the Lithium Duplex Sterling (LiDS) engine, which we assume will take 2 hr. After that ALIVE will deploy its high gain antenna and begin its Earth access routine. Once high rate communication has been established, the first 55 Mb batch of descent data will be sent to Earth, 14 Advanced Lithium Ion Venus Explorer (ALIVE) at 2 kbps. This operation will take 7.6 hr. The rest of the descent data will be sent later on bundled with the landed science data. (5) Landed Science ALIVE will toggle between periods of science data gathering (6 hr/d) and periods of data transmission back to Earth (18 hr/d). ALIVE is designed to send 130 Mb of data per day. Assuming 2 hr for the LiDS activation and 7.6 hr for the initial descent data transmission, ALIVE should have four full periods of landed science, and four full periods of data transmission. By necessity, the last science/transmission cycle will be shorter: one period of science lasting approximately 3.5 hr, followed by a transmission period of close to 11 hr. For this phase of the mission we assumed four main instruments: § Raman/Laser Induced Breakdown Spectroscopy (LIBS).—The LIBS is re-pointable by Earth command. The current design allows for 12 samples, each 12 Mb, expected total of 62.4 Mb is to be gathered. § Panoramic Camera (Pan Cam).—The Pan Cam is expected to make two eight-frame panoramas, for a total of 308 Mb of data. § Context Imager.—Expected to capture 12 images at 20 Mb each, with an expected total of 220 Mb § Meteorology Data (ASI).—Should operate at 1 bps for the duration of the science periods (27.6 hr) and a total of 100 Kb (6) End of Mission Figure 2.9 provides a graphical illustration of the ALIVE EDL operations. Figure 2.9—ALIVE EDL operations. CD–2012-72 15 March 2012 COMPASS Final Report 2.5.6 Mission Communications Details The distance between the Earth and S/C is increasing from launch until arrival. At arrival, the S/C (and Venus) are 0.7 AU apart. The Earth-Probe distance is shown in Figure 2.10(a). The Sun-Earth-Probe and Sun-Probe-Earth angles are shown in Figure 2.10(b). (a) (b) Figure 2.10—Earth-Probe distance (a) and SEP and SPE angles (b). Communication analysis during the surface stay was performed using the Satellite Orbit Analysis Program (SOAP) (Figure 2.11). § § § § § § Ovda Regio (–2.8° S, 85.6° E) was the location selected for this mission to support interesting science and increase communication opportunities with the Earth-bound DSN satellites October 24, 2023, is the primary Venus arrival date selected for the mission due to LV performance and Venus to Earth communication availability from the Ovda Regio location The ALIVE mission is currently planned to generate science data for 5 days. Communications from Ovda Regio to the Earth DSN sites is almost continuous for the 5 day period. The SOAP analysis assumes that during a communication period : − The Sun is in View from Ovda Regio and Earth − The elevation angle from the surface of Ovda Regio to Earth is > 20° − The elevation angle from the surface of Earth’s DSN’ satellites are > 20° This prevents mountainous terrain from interfering with ALIVE science At least one DSN site is in view from Ovda Regio § § The communications system was sized to account for a range of 0.74 AU (~112,000,000 km) from Ovda Regio to Earth for the 5 day mission. In the event that the mission was extended, additional opportunities would be available. 16 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 2.11—SOAP communications analysis. Figure 2.12—Need caption Figure 2.13—Need caption CD–2012-72 17 March 2012 COMPASS Final Report Figure 2.14— Figure 2.15— (This data is included when the mission trajectory will take the S/C far from the Earth and the communication system needs accurate distances.) 2.6 LV Details NASA ELV performance estimation curve(s) High energy orbits 2 2 C3 (km /s ) Figure 2.16— 18 Advanced Lithium Ion Venus Explorer (ALIVE) 2.6.1 Payload Fairing Configuration The ALIVE Lander was configured to launch atop an Atlas V 411 (performance shown in Table 2.4), inside of the 4-m LPF fairing and is required to be fully encapsulated inside an aeroshell in order to enter the Venus atmosphere. Due to encapsulation inside the aeroshell, a cruise deck is required to provide power, propulsion, and GN&C for the transit from Earth to Venus. This cruise deck will also provide the interface between the payload adaptor and aeroshell. For launch mass purposes, a C22/type D1666 Payload Adaptor (PLA) stack was assumed. Due to time constraints during the study, a CAD model of the cruise deck was not laid out. However, a cruise deck was sized by the COMPASS Team in order to obtain a mass to ensure the overall system mass fit within the LV capability as well as provide accurate mission analysis. Based on the COMPASS Team sizing, there do not appear to be any major configuration issues with the cruise deck. The aeroshell used in this design was based on the outer mold line of the aeroshell used for the Genesis mission. Both the backshell and heat shield were scaled up to obtain a maximum external diameter of 3.6m. This diameter provides sufficient volume inside the aeroshell for the Lander, and allows the aeroshell to fit within the 3.65-m diameter static envelope associated with the 4-m fairing. The overall dimensions of the aeroshell can be seen in Figure 2.17. Figure 2.17—ALIVE Lander aeroshell dimensions. In order for the ALIVE Lander to fit within the envelope of the aeroshell, several components needed to be stowed for the launch and cruise phases of the mission. These components include the three landing legs and the 0.75-m diameter X-band dish antenna and boom. The landing legs utilize a spring-lock mechanism for deployment, and are folded upwards when stowed, allowing the lower portion of the landing leg and the landing pads to fit within the envelope of the heat shield. The landing legs will be deployed just after the heat shield is jettisoned upon deployment of the parachute (stowed in the top of the backshell). The X-band antenna boom is stowed in a horizontal position, while the dish utilizes its 2-axis gimbal to position it so that it fits within the envelope of the aeroshell. Both are tied down to the large drag flap structure (discussed in Section 3.3) for launch. A single mechanism at the base of the boom is used to rotate it 90° to a vertical position upon landing on the surface of Venus. The boom is approximately 0.85-m in length, allowing the antenna to gimbal freely in two axes without any physical interference or blockage of the beam. CD–2012-72 19 March 2012 COMPASS Final Report Two isometric views of the ALIVE Lander inside the aeroshell can be seen in Figure 2.18 while the deployment sequence for the landing legs and X-band antenna can be seen in Figure 2.19. Additional images of the stowed ALIVE Lander can be found in Appendix B. Figure 2.18—Isometric views of the ALIVE Lander inside the aeroshell. Figure 2.19—Landing legs and X-band antenna deployment sequence. 3.0 BASELINE DESIGN 3.1 Top-Level Design 3.1.1 Master Equipment List (MEL) The Cruise Deck, Aeroshell, and Lander together are required to fit inside of the same physical Atlas V 411 LV along with fitting inside a total mass allocation as a requirement for this analysis. The theory behind the design of the MEL for this study is shown in Figure 3.1. The impacts of structure, performance, and thermal are common to the elements of the ALIVE S/C. 20 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 3.1—ALIVE design approach. Therefore, the MEL lists these three major elements in terms of the major subsystems within them. The ALIVE S/C, previously named the Extended Venus Explorer (EVE), is listed as work breakdown structure (WBS) Element 06. The Lander itself is listed in the MEL as WBS Element 06.1. The Aeroshell, is listed as WBS Element 06.2, and the Cruise Deck is listed as WBS Element 06.3 respectively. Table 3.1 shows the MEL listing of the Lander, Aeroshell, and Cruise Deck as the three elements of the ALIVE S/C designed by the COMPASS Team and documented in this study. Table 3.1—ALIVE MEL WBS Format WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.1 06.1.2 06.1.3 06.1.4 06.1.5 06.1.6 06.1.11 06.2 06.2.2 06.2.4 06.2.6 06.2.11 06.3 06.3.2 06.3.3 06.3.5 06.3.6 06.3.7 06.3.8 06.3.11 QTY (kg) Lander Science Attitude Determination and Control Command & Data Handling Communications and Tracking Electrical Power Subsystem Thermal Control (Non-Propellant) Structures and Mechanisms Aeroshell Attitude Determination and Control Communications and Tracking Thermal Control (Non-Propellant) Structures and Mechanisms Cruise Deck Attitude Determination and Control Command & Data Handling Electrical Power Subsystem Thermal Control (Non-Propellant) Propulsion (Chemical Hardware) Propellant (Chemical) Structures and Mechanisms CD–2012-72 Unit Mass 21 Basic Mass Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 39.80 18.7% 7.45 47.25 142.61 17.4% 24.75 167.36 22.60 33.0% 7.47 30.07 47.71 10.9% 5.20 52.91 277.50 12.2% 33.77 311.27 35.79 18.0% 6.44 42.23 513.91 18.0% 92.50 606.42 608.77 18.0% 109.47 718.24 54.23 18.0% 9.76 63.99 1.40 10.0% 0.14 1.54 371.29 18.0% 66.83 438.13 181.85 18.0% 32.73 214.58 229.25 9.4% 21.44 250.69 3.44 3.0% 0.10 3.54 7.50 14.0% 1.05 8.55 33.00 3.0% 1.00 34.00 10.34 18.0% 1.86 12.20 30.52 5.2% 1.58 32.10 56.43 0.0% 0.00 56.43 88.01 18.0% 15.84 103.86 March 2012 COMPASS Final Report The Lander, Aeroshell, and Cruise Deck sections of the MEL starts at WBS 06.1, WBS 06.2, WBS 06.3, and opens down to the subsystem level, as shown in Table 3.2. The Lander science instruments can be found within WBS 06.1.1, and discussed in Section 5.1. 3.1.2 S/C Total Mass Summary The system-level summary for the baseline case, which includes the additional system-level growth, is shown in Table 3.2. In order to reach the 30% total system level growth on the basic mass of the S/C required for this study, MGA and system level growth was calculated for each individual subsystem within the three elements. Table 3.2—ALIVE System Summary Basic Mass (kg) WBS Main Subsystems 06 Extended Venus Explorer (EVE) Spacecraft 06.1 06.1.1 06.1.2 06.1.3 06.1.4 06.1.5 06.1.6 06.1.7 06.1.8 06.1.9 06.1.10 Lander Science Attitude Determination and Control Command and Data Handling Communications and Tracking Electrical Power Subsystem Thermal Control (Non-Propellant) Propulsion (Chemical Hardware) Propellant (Chemical) Propulsion EP Hardware) Propellant (EP) Structures and Mechanisms 06.2 06.2.2 06.2.3 06.2.4 06.2.5 06.2.6 06.2.7 06.2.8 06.2.9 06.2.10 06.3 06.3.1 06.3.2 06.3.3 06.3.4 06.3.5 06.3.6 06.3.7 06.3.8 06.3.9 06.3.10 06.3.11 2226.4 1257.5 16% 39.8 7.4 47.2 19% 142.6 24.7 167.4 17% 22.6 7.5 30.1 33% 11% 47.7 5.2 52.9 277.5 33.8 311.3 12% 35.8 6.4 42.2 18% 0.0 0.0 0.0 0.0 0.0 513.9 92.5 606.4 1080 1404 1080 324 146 324 608.8 109.5 718.2 0.0 0.0 0.0 54.2 9.8 64.0 0.0 0.0 0.0 18% Total Growth 30% 14% 1404 1.4 0.1 1.5 0.0 0.0 0.0 371.3 66.8 438.1 0.0 0.0 0.0 0.0 18% 18% 10% 18% 0.0 0.0 0.0 Cruise Deck Science Attitude Determination and Control Command and Data Handling Communications and Tracking Electrical Power Subsystem Thermal Control (Non-Propellant) Propulsion (Chemical Hardware) Propellant (Chemical) Propulsion EP Hardware) Propellant (EP) Structures and Mechanisms 0.0 0.0 0.0 0.0 0.0 181.8 32.7 214.6 609 183 73 183 791 609 229.2 21.4 250.7 0.0 0.0 0.0 18% Total Growth 30% 12% 791 9% 3.4 0.1 3.5 3% 7.5 1.1 8.6 14% 0.0 0.0 0.0 33.0 1.0 34.0 3% 10.3 1.9 12.2 18% 30.5 1.6 32.1 5% 56.4 0.0 56.4 0.0 0.0 88.0 System LeveL Growth Calculations_Cruise Deck Dry Mass Desired System Level Growth Additional Growth (carried at system level) Total Wet Mass with Growth § 308.5 177.6 0.0 06.2.11 System LeveL Growth Calculations _ Aeroshell Dry Mass Desired System Level Growth Additional Growth (carried at system level) Total Wet Mass with Growth Aggregate Growth (%) 1917.9 0.0 Aeroshell Science Attitude Determination and Control Command and Data Handling Communications and Tracking Electrical Power Subsystem Thermal Control (Non-Propellant) Propulsion (Chemical Hardware) Propellant (Chemical) Propulsion EP Hardware) Propellant (EP) Structures and Mechanisms Total Mass (kg) 1079.9 0.0 06.1.11 System LeveL Growth Calculations _Lander Dry Mass Desired System Level Growth Additional Growth (carried at system level) Total Wet Mass with Growth 06.2.1 Growth (kg) 173 229 0.0 0.0 15.8 52 30 52 103.9 18% 225 Total Growth 30% 18% 281 The Lander MGA was 16%, and the remaining 14% growth (146 kg) was carried at the system level 22 Advanced Lithium Ion Venus Explorer (ALIVE) § § The Aeroshell MGA was 18%, and the remaining 12% growth (73 kg) was carried at the system level The Cruise Deck (contains RCS) MGA was 9%, with the remaining 21% growth (30 kg) carried at the system level This additional system-level mass is counted as part of the inert mass to be flown along the required trajectory. Therefore, the additional system-level growth mass impacts the total propellant required for the mission design. The total wet mass of the ALIVE S/C stack with system level growth and MGA (558 kg) included was 2476 kg. Sections 3.2.1 and 3.2.2 provide additional details about the basic and total masses of the different subsystems and the entire ALIVE S/C, after MGA and system level growth has been applied. In the calculations shown in Table 3.3, the inert mass of the ALIVE S/C is the dry mass plus trapped pressurant, residuals, and propellant margin. The dry mass on each segment is calculated as the total bottoms-up dry mass with the MGA percentage applied plus additional system mass, so that the total growth on each stage is 30% of the basic mass. The total dry basic mass of the ALIVE S/C Stack is 1862 kg. The total basic mass of the ALIVE S/C with the bottoms-up growth (308 kg of the dry mass applied by the subsystem engineers) is 1862 kg + 308 kg = 2170 kg. This is also known as predicted mass, and does not contain the system level growth to reach the 30% growth on dry mass. The total inert mass of the ALIVE S/C with 30% growth carried on the basic masses is 2427 kg. The total wet mass of the complete ALIVE stack is 1918 kg + 558 kg = 2476 kg. This summary of mass is shown in Table 3.3. Table 3.3—ALIVE Total Mass With Payload (Includes 30% System Level Growth) Total masses Total stack dry ..............................................................2420 kg Total stack inert ............................................................2427 kg Total stack wet ..............................................................2476 kg Total Lander dry............................................................1404 kg Total Lander inert..........................................................1404 kg Total Lander wet ...........................................................1404 kg Total Aeroshell dry ..........................................................791 kg Total Aeroshell inert ........................................................791 kg Total Aeroshell wet .........................................................791 kg Total Cruise Deck dry .....................................................225 kg Total Cruise Deck inert ...................................................232 kg Total Cruise Deck wet.....................................................281 kg 3.1.3 Power Equipment List (PEL) Table 3.4—Definition of the ALIVE S/C Power Modes Mode Power mode 1 Power mode 2 Power mode 3 Power mode 4 Power mode 5 Power mode 6 Title Ground Ops & Launch SA Deploy & Cruise Descent Drop Cruise Deck, Heat shield, Parachute, Aeroshell Free Fall Descent Landed Science Mode Landed Communication Mode Description Preliminary ground operations, transfer to internal power, launch, and insertion into Venus trajectory Deployment of solar arrays, ALIVE generating power, and transit to Venus Entry, drop of cruise deck, heat shield, deployment of parachute, parachute release, aeroshell release, and first part of descent science Free fall portion of descent and descent science Portion of the mission devoted to gathering science Portion of the mission devoted to communication Table 3.5 provides the assumptions about the power requirements in all the modes of operation. The power system designers use these assumptions to size the solar arrays and other power system components. CD–2012-72 23 March 2012 COMPASS Final Report Table 3.5—ALIVE S/C PEL Table 3.6 shows the thermal waste heat for the ALIVE S/C. The thermal waste heat data is used by the Thermal subsystem lead to size each of the ALIVE elements for worst-case environmental conditions. Table 3.6—Case x Thermal Waste Heat Per Power Mode 24 Advanced Lithium Ion Venus Explorer (ALIVE) 3.2 System-Level Summary (Needs to be updated-ces) The system block diagram that captures the theory behind the ALIVE design is shown in Figure 5.1. The components were designed and placed in a manner that allows for a controlled landing at Ovda Regio while supporting descent and surface science. 3.2.1 Propellant Calculations The propellant details are captured in Table 3.7. The total 2476 kg stack wet mass includes residuals and margin from each of the three elements. The total 2427 kg S/C inert mass is used by the mission seat to iteratively calculate total useable propellant. Table 3.7—ALIVE S/C Propellant Details Lander: Propellant Details (Chemical) Lander Totals Lander Dry mass ........................................................................ 1404 kg Lander Inert mass ...................................................................... 1404 kg Lander Wet mass ....................................................................... 1404 kg Aeroshell: Propellant Details (Chemical) Aeroshell Totals Aeroshell Dry mass ...................................................................... 791 kg Aeroshell Inert mass .................................................................... 791 kg Aeroshell Wet mass ..................................................................... 791 kg Cruise Deck: Propellant Details (Chemical) RCS/ACS Used Prop ..................................................................... 49 kg Mass, RCS Total ............................................................................ 56 kg RCS/ACS margin ............................................................................. 5 kg RCS/ACS Residuals ........................................................................ 2 kg RCS Total Loaded Pressurant ......................................................... 1 kg Cruise Deck Totals Cruise Deck Dry mass ................................................................. 225 kg Cruise Deck Inert mass ................................................................ 232 kg Cruise Deck Wet mass................................................................. 281 kg The formulas given below were used to calculate the amount of propellant needed to push the ALIVE S/C (Lander, Aeroshell, and Cruise Deck) along the trajectory to the surface of Venus. The used propellant is calculated using the following rocket equation: ⎛ m ⎞ ΔV = Isp * g * ln⎜⎜ 0 ⎟⎟ ⎝ m1 ⎠ which can be rewritten as: ⎛ − ΔV ⎞ ⎟⎟ ⎜⎜ m1 = m0 * e ⎝ Isp*g ⎠ The variables in this equation are signified as follows: ∆V is the total mission change in velocity to perform the attitude control maneuvers m0 is the initial total mass, including propellant m1 is the final total mass and is the value being determined, as shown by the second equation Isp is the specific impulse expressed as a time period g is the gravitational constant, which is equal to 9.8 m/s Following are propellant details for the mission. Additional information can be found in Table 3.7. § § Total RCS/ACS propellant = (Used + Margin + Residuals + Loaded Pressurant) = 49 kg + 5 kg + 2 kg + 1kg = 56 kg Total ALIVE Stack Masses: CD–2012-72 25 March 2012 COMPASS Final Report − Wet mass = (basic mass + subsystem growth + system growth + total propellant + total RCS propellant) = 2476 kg − Dry mass = (wet mass – total propellant) = 2420 kg − Inert Mass = (wet mass – used propellant) = 2427 kg Table 3.8—Inert Mass Calculations For ALIVE Total S/C ALIVE S/C Mass Calculations EVE spacecraft Total Wet Mass EVE spacecraft total Dry Mass Dry Mass Desired System Level Growth Additional Growth (carried at system level) Tot al Useable Propellant Tot al Trapped Propellants, Margin, pressurant Tot al Inert Mass with Growth EVE spacecraft Total Wet Mass with Sy stem Level Growth Basic Mass (kg) Growth (kg) To tal Mass (kg) 1918 1862 1862 308 308 558 250 2226 2170 2420 49 7 1869 1918 558 558 Aggregate Growth (%) 16% 30% 14% 49 7 2427 2476 The LV performance margin of 584 kg was calculated by subtracting the wet mass of the S/C from the assumed LV performance. After including an additional margin of 10% from the LV performance, the ALIVE S/C was required to be lighter than 3060 kg. Table 3.9—ALIVE Architecture Details Architecture Details LV .......................................................................................... Atlas V 411 V .............................................................................................. 2.65 km/s 2 2 Energy, C3 ............................................................................. 7.00 km /s ELV performance (pre-margin) ...................................................3400 kg ELV Margin (%) ................................................................................ 10% ELV performance (post-margin) ..................................................3060 kg C22 ELV Adaptor (Stays with ELV) ..................................................0 kg ELV performance (post-adaptor) ................................................3060 kg EV Spacecraft Total Wet Mass with System Level Growth ........2476 kg Available ELV Margin ................................................................ 584 kg Available ELV Margin (%) .............................................................. 19% Lander: Propellant Details (Chemical) Lander Dry mass .........................................................................1404 kg Lander Inert mass .......................................................................1404 kg Lander Wet mass ........................................................................1404 kg Aeroshell: Propellant Details (Chemical) Aeroshell Dry mass .......................................................................791 kg Aeroshell Inert mass .....................................................................791 kg Aeroshell Wet mass ......................................................................791 kg Cruise Deck: Propellant Details (Chemical) RCS/ACS Used Prop ......................................................................49 kg Mass, RCS Total .............................................................................56 kg RCS/ACS margin ..............................................................................5 kg RCS/ACS Residuals .........................................................................2 kg RCS Total Loaded Pressurant ..........................................................1 kg Cruise Deck Totals Cruise Deck Dry mass ..................................................................225 kg Cruise Deck Inert mass .................................................................232 kg Cruise Deck Wet mass .................................................................281 kg ∞ The mass of the ELV is absorbed in the structure calculations. 26 Advanced Lithium Ion Venus Explorer (ALIVE) 4.0 AREAS FOR FUTURE STUDY The ALIVE landed duration is only limited by the amount of Li which can be carried by the lander. Further studies are needed to investigate how additional mass and volume of Li can be carried, in the minimum by a more elegant Li tank design perhaps even longer using a larger launcher and/or larger aeroshell. Other power conversion/cooling systems might also bring other benefits. A more detailed conceptual design of the Li burner system is necessary for technology development planning purposes. 5.0 SUBSYSTEM BREAKDOWN 5.1 Science Package 5.1.1 Descent Instruments The ALIVE science package consisted of various descent and surface science instruments, see Table 5.1 and Table 5.3. Table 5.1—Descent Instruments Instrument NMS Mass (kg) 11 4.5 TLS Descent imager ASI IMU 2 2 ----- Power (W) 50 Footprint (m) 0.26 by 0.16 by 0.39 Data (kbps) 0.5 17 0.25 by 0.10 by 0.10 1.0 12 3.2 ----- 0.15 by 0.15 by 0.10 0.10 by 0.10 by 0.10 ------------------------- 24 0.25 0.5 Heritage High: MSL, SAM, Pioneer High: MSL, SAM High: MSL High: flagship High comments A slightly smaller instrument was flown on Pioneer Venus Data rate can be reduced (will give fewer points in profile) Only used last 10 km of descent Data rate seems to be high Assume MEMS accelerometer IMU The 3-axis accelerometer (IMU) is part of the atmospheric science, to measure wind velocities from descent motion. Table 5.1 does not include IMU mass or power because the IMU instrument is accounted in the G&NC budget. ASI This consists primarily of temperature and pressure measurements during descent. Ten 12-b measurements per second should be sufficient, that would be 0.12 kbps. If we run the anemometer during descent; this will double the bit rate. The data rate from the VITAL statistics is 2.5 kbps; this seems higher than is needed. Descent imager data rate: The images are assumed to begin at 10 km, and the descent rate is assumed to be 5m/sec, so the duration is 2000 s. The ten lossless images (48 Mb) is thus an average rate of 24 kbps. Data rate will be lower if we assume a lower descent rate or higher data compression. Since the highest altitude frames will be blurred due to atmospheric scattering, it may be reasonable to use higher compression for all but the lowest few frames Data Volume § NMS data volume calculation: − Assume one measurement every 100 m from 30 km to surface = 300 measurements. − Each measurement is 12 b times 512 data points = 6 Kb (512 data points will give 0.2 Dalton resolution for 1 to 99 Dalton range. This is comparable to Cassini data resolution) CD–2012-72 27 March 2012 COMPASS Final Report − Total is 1.8 Mb − If these measurements are taken over a descent time of 1 hr (3600 s), data rate is 0.5 Kb/s Cassini instrument: http://lasp.colorado.edu/~horanyi/graduate_seminar/Ion_Neutral_Mass_Spec.pdf TLS data volume: − § 5.1.2 − Assume one measurement every 100 m from 30 km to surface = 300 measurements. − Each measurement is 12 b times 1024 data points = 12 Kb − Total is 3.6 Mb − If these measurements are taken over a descent time of 1 hr (3600 s), this will come to 1 Kb/s Surface Instrument Details Table 5.2—Surface Instruments Instrument LIBS Mass (kg) 13 Pan Cam 1 Power (W) 50 2.2 Context Imager 2 2.2 Meteorology (ASI) 0.1 3.2 Footprint (m) Data Heritage comments (Mb) 5.2/sample Will be “12 b, three measurements per demonstrated sample” (1 R, 2 LIBS) on MSL Two boxes (laser and spectrometer): 0.15 by 0.15 by 0.30 0.20 by 0.20 by 0.20 Two boxes (optical and 154 total electronics) 0.04 by 0.05 by 0.06 0.07 by 0.07 by 0.034 Two boxes (optical and 20/image electronics) 0.04 by 0.05 by 0.06 0.07 by 0.07 by 0.034 0.05 by 0.05 by 0.15 1 bps High: MSL Data rate can be reduced with higher compression if needed. Mass includes window High: MSL Data rate can be reduced with higher compression if needed. Mass includes window High: flagship Mass includes only Anemometer LIBS/Raman The LIBs instrument has an optical head with the laser and mirror, and a separate spectrometer connected to the optical head with a fiber optic. The mirror diameter for the MSL instrument was 11 cm; the larger the mirror, the farther away the instrument can take measurements. For a baseline, we need a window with an 11 cm diameter at the outside (the window can be a truncated cone that tapers to a smaller size on the inside). The LIBS will have an externally-mounted mirror that uses high-temperature motors to adjust the pointing in two axes. For info and photos of the MSL instrument, see http://www.nasa.gov/mission_pages/msl/multimedia/gallery/pia13398.html and http://mslscicorner.jpl.nasa.gov/Instruments/ChemCam/ Panoramic Imager The panoramic imager has a separate window, and also is pointed using an externally-mounted mirror Meteorology Meteorology measurements will include the temperature and pressure sensors from the descent ASI package. The instruments are already incorporated into the descent instrument list, and hence only the anemometer mass and volume is included here. The Anemometer is a rod that will protrude 15 cm upwards from the lander. Data volume for Images 28 Advanced Lithium Ion Venus Explorer (ALIVE) Calculation of data volume for the images: Compression: The MER Pan Cam investigation did lossless (“LOCO”) compression at 4.8 bits per pixel (bpp). We can probably do better than this, however, this value will be used for calculations. Descent imager data rate assumption: The science minimum is assumed to be acquisition of ten 1024 by 1024-pixel frames. These will be compressed using LOCO at 4.8 bpp. The total data volume is thus 48 Mb. Panorama: The field of view is 60°; we need some overlap to make a panorama, and so the full panorama requires eight frames. Each frame is 2048 by 2048 pixels = 4 Megapixels. We will take the color image in two parts, a lossless black and white image, and then a higher compression for the four frames of color (the color frames are not going to very different from the black and white, so this can be highly compressed with no loss of image quality). The black and white panorama is thus (eight images) times (4 M-pixels/image) times (4.8 bpp) = 154 Mb. The color portion of the data will be encoded to 1 bpp per color. The color data for the panorama is thus (eight frames) times (four colors per frame) times (4 M-pixels/image) times (1 bpp) = 118 Mb. 5.1.3 Science Design and MEL The full science payload, summarized in the MEL for the ALIVE S/C in Table 5.3, consists of the descent science instruments, surface science instruments, and additional instruments on the Lander. Table 5.3—Science ALIVE MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design (kg) Lander Science 06.1 06.1.1 Descent Science Instruments Neutral Mass Spectrometer (NMS) Tunable Laser Spectrometer (TLS) Descent imager Atmospheric structure (ASI) Surface Science Instruments Raman / Laser Induced Breakdown Specroscopy (LIBS) Box 1 Panoramic Imager Optical Box Context Imager Optical Box Meteorology (ASI) Raman / Laser Induced Breakdown Specroscopy (LIBS) Box 2 Additional Instruments Motors for Pointing Optical Instruments Panoramic Imager Electronics Box Context Imager Electronics Box 06.1.1.a 06.1.1.a.a 06.1.1.a.b 06.1.1.a.c 06.1.1.a.d 06.1.1.b 06.1.1.b.a 06.1.1.b.b 06.1.1.b.c 06.1.1.b.d 06.1.1.b.e 06.1.1.c 06.1.1.c.a 06.1.1.c.b 06.1.1.c.c 5.2 Communications 5.2.1 Communications Requirements § QTY Unit Mass Basic Mass Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 39.80 18.7% 7.45 47.25 19.50 20.0% 3.90 23.40 1 11.00 11.00 20.0% 2.20 13.20 1 4.5 4.50 20.0% 0.90 5.40 1 2.0 2.00 20.0% 0.40 2.40 1 2.0 2.00 20.0% 0.40 2.40 3.02 18.12 15.10 20.0% 1 6.5 6.50 20.0% 1.30 7.80 2 0.5 1.00 20.0% 0.20 1.20 1 1.0 1.00 20.0% 0.20 1.20 1 0.1 0.10 20.0% 0.02 0.12 1 6.5 6.50 20.0% 1.30 7.80 5.20 10.2% 0.53 5.73 4 0.80 3.20 4.0% 0.13 3.33 2 0.50 1.00 20.0% 0.20 1.20 1 1.00 1.00 20.0% 0.20 1.20 Communications design philosophy − CD–2012-72 Provide direct to Earth communication during all phases of operation 29 March 2012 COMPASS Final Report − Provide the highest possible data rates for science. Target 2.2 kbps. − Single fault tolerant − Flight heritage components − Low power consumption electronics, except RF transmitter − Single Event Upset (SEU) tolerant electronics − Software hard coded into ASICS chips − Use of DSN antenna arraying capabilities for increase receive aperture − X-Band was directed for communications The communications link budget for the ALIVE S/C can be found in Table 5.4. 5.2.2 Communications Assumptions Hardware Functionality § § § § § § § Antennas two fly away low gain antennas (LGA), two LGA’s on Lander and a high gain antenna (HGA) for primary landed communications. LGA designed by Allan Hanson, Hughes Aircraft Company for Venus probe (Figure 5.2) HGA includes deployment mechanisms, two access gimbals and rotary joints (Figure 5.3) HGA a special RF waveguide/window to pierce shell of Lander for reduced heat transference ~ 4 wavelengths depth Software functionality − Embedded software, vender specific language Primary communications mass: 47 kg Design based on current hardware: LRO and Orion HGA’s, the Deep Space Transponder and currently deployed TWTA’s by Boeing (Figure 5.5) Table 5.4—Communications Science Link Budget Transmitter Transmitter power (W, dBW) Losses of antenna (dB) Efficiency Transmitted power (W, dBW) DC power 75 W ---------------------0.5 59.57 W 150 W Transmit antenna Frequency 8.4 GHz Dish diameter 0.75 m Directivity 4358.52 Antenna efficiency 0.5 Antenna gain 2179.26 EIRP (dBW) ---------------------Receiver ---------------------Receiver noise figure ---------------------Receiver noise temperature ---------------------Receiver antenna diameter 70 m Directivity 37967522.42 Antenna efficiency 0.63 Antenna gain 23919539.12 ---------------------Distance between antennas 114000000 km –29 Spreading loss 3.88×10 Receiver noise temperature (k)/noise figure (dB) 81.52 K Bandwidth (Hz) 4000 –18 Spectral power density 4.50×10 W 30 18.75 dBW –1 dBW ---------------------17.75 dBW ---------------------------------------------------------------36.39 dBi ---------------------33.38 dBi 51.dBW ---------------------1.0 dB 81.52 K ---------------------75.79 dB ---------------------73.79 dB 56 ---------------------–284.11 dB 1.1 dB ---------------------–173.47 dBW Advanced Lithium Ion Venus Explorer (ALIVE) Bits per Hz SNR Eb/No 10*log(2) = 3.01 db Qpsk = 2 Required SNR Es/No –7 Margin 0.55 48.65 ---------------------3.01 ---------------------------------------------------------------- ---------------------16.87 dB ------------------------------------------2.189291851 dB ---------------------14.68 dB Figure 5.1—Block diagram of ALIVE communications hardware – based on Venus Probe CD–2012-72 31 March 2012 COMPASS Final Report Figure 5.2—Illustration of Venus Probe LGA Figure 5.3—Graphic of Orion 32 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 5.4—Image of Representive SDST communications hardware. Figure 5.5—Image of Representative TWTA and EPC. 5.2.3 Communications Design and MEL Table 5.5—Communications Case 1 MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.4 (kg) X Band System SDT Transponder X Band gimbaling antenna X Band antenna Wave guide X Band TWTA and EPC X Band LNA Low Gain Antenna SC positive Low Gain Antenna SC negative Diplexer Switch A Switch B Communications Instrumentation Cables TPS 06.1.4.a.a 06.1.4.a.b 06.1.4.a.c 06.1.4.a.d 06.1.4.a.e 06.1.4.a.f 06.1.4.a.g 06.1.4.a.h 06.1.4.a.j 06.1.4.a.k 06.1.4.a.l 06.1.4.e 06.1.4.e.b 06.1.4.e.c 06.2.4 Unit Mass Lander Communications and Tracking 06.1.4.a 06.2 QTY 2 3.20 1 18.00 1 1 06.2.4.a.a CD–2012-72 Total Mass (kg) (%) (kg) (kg) 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 47.71 10.9% 5.20 52.91 40.15 12.9% 5.20 45.35 10.0% 0.64 7.04 18.00 10.0% 1.80 19.80 1.45 1.45 10.0% 0.15 1.60 0.50 0.50 30.0% 0.15 0.65 2 3.70 7.40 10.0% 0.74 8.14 2 0.70 1.40 30.0% 0.42 1.82 1 0.50 0.50 10.0% 0.05 0.55 1 0.50 0.50 10.0% 0.05 0.55 2 0.50 1.00 30.0% 0.30 1.30 1 1.50 1.50 30.0% 0.45 1.95 1 1.50 1.50 30.0% 0.45 1.95 1 3.78 1 3.78 2 33 Growth 1917.94 7.56 X Band System LGA Fly Away postive and negative Growth 6.40 Aeroshell Communications and Tracking 06.2.4.a Basic Mass 0.70 3.78 3.78 0.0% 0.0% 0.0% 608.77 18.0% 1.40 1.40 1.40 0.00 7.56 0.00 0.00 3.78 3.78 109.47 718.24 10.0% 0.14 1.54 10.0% 0.14 10.0% 1.54 0.14 1.54 March 2012 COMPASS Final Report 5.2.4 Communications Recommendation Development of high temperature electronics to make possible an X-Band phased array. Research of propagation loss in the Venus atmosphere at the assigned frequency may increase the probability of returning all mission data. 5.3 Command and Data Handling The main purpose of the C&DH system is collecting and distributing non-flight-critical sensor data from the instrumentation throughout the mission and storing it in local memory via high-speed data buses. GN&C, propulsion, and thermal control requirements indicate the need for controlling valves and gimbals, as well as sensing pressure and temperature transducers. All telemetry acquisition and processing of data is followed by forwarding the data to the communication subsystem for transmission to Earth. 5.3.1 C&DH Requirements The design requirements for the C&DH system are as follows: § § § § § 5.3.2 Avionics components and parts shall be Class S, per MIL–STD–883B. Avionics shall be one fault tolerant using cold spares. Data storage unit shall provide at least 5 GB of onboard permanent solid-state memory. Avionics shall be ground-bonded and surge-protected to resist on-pad lightning damage. Avionics shall be designed to withstand the on-orbit ionizing and non-ionizing radiation environments dictated by the mission profile. It is important to avoid over-specifying the radtolerance levels to minimize cost for parts and testing. C&DH Assumptions The following design assumptions are based on the mission requirements: § § § Implemented with rad-tolerant microcontrollers, field-programmable gate arrays (FPGAs), and data storage using solid-state random access memory (RAM) and Flash memory. The LEON3 processor is an example of a modern rad-tolerant microcontroller. Avionics spare circuitry for fault tolerance is implemented as cold spares in order to minimize power consumption. Hardware design heritage is based on previous S/C and lessons learned. − 5.3.3 Sensor estimate is based on a preliminary assumption of number of channels for input and output and likely will decrease as the design stabilizes. C&DH Design and MEL The C&DH system consists of 100 MIPS LEON3-class processor boards containing various hardware and software mechanisms such as timeouts and watchdog circuitry to provide for single fault tolerance. Each processor board includes an FPGA-embedded core built with a main processor such as the LEON3 series, capable of supporting C&DH functions, a 5-plus GB solid-state memory card, as well as communications and payload interface cards. The primary processor is capable of autonomous failover to a redundant cold spare unit if a fault is detected. Depending on choice of processor, flight computers will use a real-time operating system such as VxWorks or Green Hills Integrity. To support all mission phases, the number of source lines of code (SLOC) has been estimated to be 250000 SLOCs. However, this estimate and implied development cost should be tempered with the understanding that recent developments in autocode technologies that generate known good instruction loads will become a design standard. 34 Advanced Lithium Ion Venus Explorer (ALIVE) The following list is comprised of the main avionics components and their quantities, as input to the MEL shown in Table 5.6: § § § § § Main computers (one main computer and one redundant cold spare) Data acquisition channels (including redundant paths for single-fault tolerance) Cruise Deck has a simple DCIU commanded by the Lander Redundant solid-state memory Instrumentation (including approximately 40 sensors, mass of 6 ounces each, power requirement of 50 mW each) Note: As shown in the MEL, the initial estimate contained two 48-channel analog-to-digital and digitalto-analog serial digital interface (SDI) cards and one 48-channel serial data output (SDO) card, giving 144 channels of input/output, not including any serial bus input/output, all used to estimate worst-case mass and power. § S/C cabling (per Monte Carlo simulation): − Instrumentation wiring approximately 11m per sensor run − Approximately 583 m total, 20-24 American Wire Gauge (AWG) Tefzel (exclusive of high currents Power system conductors) Table 5.6—C&DH ALIVE S/C MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.3 (kg) C&DH Hardware FPGA IP CPU rad hard LEON3 Watchdog switcher Time Generation Unit Mass Memory Module Command and Control Harness cPCI enclosure with power supply Valve drivers Igniter drivers Separation drivers TVC drivers SLOCS Instrumentation & Wiring AD/DA/SDI card SDO card Pressure and Temperature Sensors 06.1.3.a.a 06.1.3.a.b 06.1.3.a.c 06.1.3.a.d 06.1.3.a.e 06.1.3.a.f 06.1.3.a.g 06.1.3.a.h 06.1.3.a.i 06.1.3.a.j 06.1.3.a.m 06.1.3.b 06.1.3.b.a 06.1.3.b.c 06.1.3.b.d 06.3.3 Unit Mass Lander Command & Data Handling 06.1.3.a 06.3 QTY 06.3.3.a.b 06.3.3.a.k CD–2012-72 35 Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 22.60 33.0% 7.47 30.07 19.30 33.5% 6.48 25.78 1.50 3.00 30.0% 0.90 3.90 1 0.50 0.50 30.0% 0.15 0.65 1 0.50 0.50 3.0% 0.02 0.52 1 0.50 0.50 30.0% 0.15 0.65 1 6.60 6.60 50.0% 3.30 9.90 1 5.00 5.00 20.0% 1.00 6.00 1 0.80 0.80 30.0% 0.24 1.04 1 0.80 0.80 30.0% 0.24 1.04 1 0.80 0.80 30.0% 0.24 1.04 1 0.80 0.80 30.0% 0.24 1.04 250000 0.00 0.00 0.0% 0.00 3.30 C&DH Hardware DCIU Harness Growth 2 30.0% 0.99 0.00 4.29 1 1.00 1.00 30.0% 0.30 1.30 1 1.30 1.30 30.0% 0.39 1.69 20 0.05 1.00 30.0% 0.30 Cruise Deck Command & Data Handling 06.3.3.a Basic Mass 229.25 9.4% 7.50 7.50 1.30 21.44 250.69 14.0% 1.05 8.55 14.0% 1.05 8.55 1 3.50 3.50 30.0% 1.05 4.55 1 4.00 4.00 0.0% 0.00 4.00 March 2012 COMPASS Final Report 5.3.3.1 Flight Computers and Software The flight computers and software provide the following functions: § § § § § § § § § § 5.3.4 Load, initialization, executive functions, and utilities executed by the processors Flight computer board redundancy management Data acquisition and control Command and telemetry processing via RS-422 or SERDES Health monitoring and management Power management, control, and distribution GN&C calculations Ephemeris calculations for available data communications with Earth Event sequence management Fault detection, diagnostics, and recovery C&DH Trades The S/C must have sufficient particle shielding for the avionics to withstand long-term deep-space exposure to heavy ions. Therefore, future studies should consider trading the inclusion of additional particle shielding in the avionics enclosures. In some cased, titanium (Ti) instead of aluminum (Al) can be used to add shielding with less mass due the barns ratio of Ti to Al. By mid-decade, advances in semi-automatic code generation will help guarantee a very capable, secure, and reliable operating system execution. Therefore, the choice of which computer operating system to include on a S/C designed for 2020 and beyond may not be the correct one for a S/C designed in 2012 to 2014. A final choice of operating system should await the actual beginning of detailed design. 5.3.5 C&DH Analytical Methods As a matter of common practice, the design of a new S/C’s C&DH system is often based on one that is proven effective (high TRL) on another S/C, and that requires minor or no modifications for the mission currently under development. This C&DH system is based on previous S/C, such as Dawn, New Horizons, and Extrasolar Planet Observation (EPOXI). 5.3.6 C&DH Risk Inputs C&DH risks include the following: § § § § 5.3.7 Particle radiation Launch vibration stresses Obsolescence and/or availability of low-volume space-qualified EEE parts Inability to accurately define design and performance requirements and margins early in the project, thereby leading to a system design that is unable to meet downstream requirements leading to schedule delays and cost overruns. C&DH Recommendation The following are the recommendations of the C&DH subsystem lead: § § The S/C must have sufficient EMI/RFI shielding as well as being sufficiently ground-bonded and surge-protected to resist on-pad lightning damage. The S/C must have sufficient electromagnetic/radio frequency interference and particle shielding, due to its long-term space orbital time. 36 Advanced Lithium Ion Venus Explorer (ALIVE) § Long-term availability and reliability of Avionics for the length of this mission is crucial for mission success. 5.4 Guidance, Navigation and Control 5.4.1 GN&C Requirements The GN&C subsystem is required to provide attitude determination and control throughout the entire mission, including post LV separation, cruise to Venus, and EDL. The GN&C subsystem is also required to provide an EDL profile where the vehicle experiences no more than a 40 g load. 5.4.2 GN&C Assumptions Parachute design: § Consists of determining the required canopy area and estimating the mass of the parachute § Bridle and suspension line length are left for future work Atmospheric entry defined as: § Altitude = 200 km § Velocity = 11.3 km/s 5.4.3 GN&C Design and MEL Cruise deck The GN&C hardware on the cruise deck consists of two Technical University of Denmark (DTU) Advanced Stellar Compass (ASC) Star Trackers and eight sun sensors. A single ASC data processing unit (DPU) is capable of processing information from two optical units (OU) however to remain single fault tolerant, two DPUs were employed, resulting in two DPUs and two OUs. The sun sensors provide rough attitude determination as well as knowledge of the direction to the sun during any required safe modes. Aeroshell The only GN&C hardware located in the aeroshell is the parachute. The parachute was sized to create a sufficient difference in drag acceleration between the lander and the heat shield so as to ensure no recontact by the heat shield when it gets jettisoned. Lander: The lander GN&C hardware consists of one internally redundant Northrop Grumman Scalable Inertial Measurement Unit (SIRU) that provides knowledge of vehicle body rates, position and attitude information between navigation updates, and knowledge of vehicle accelerations. Even though the SIRU is located in the lander, it provides this information during cruise as well as during EDL. In addition to the SIRU, the lander also contains the drag flap, which provides drag on the vehicle during the last phase of descent to reduce the vehicle terminal velocity. A summary of the GN&C MEL for ALIVE can be seen in Table 5.4. Table 5.7—GN&C ALIVE S/C MEL CD–2012-72 37 March 2012 COMPASS Final Report WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.2 06.1.2.a 06.1.2.a.a 06.1.2.a.b 06.2 06.2.2 06.2.2.a 06.2.2.a.c 06.3 06.3.2 06.3.2.a 06.3.2.a.b 06.3.2.a.d 06.3.2.a.e QTY Unit Mass (kg) Lander Attitude Determination and Control Guidance, Navigation, & Control Inertial Measurement Units Drag Flaps 1 7.10 1 135.51 Aeroshell Attitude Determination and Control Guidance, Navigation, & Control Main Parachute 1 54.23 Cruise Deck Attitude Determination and Control Guidance, Navigation, & Control Sun Sensors Star Tracker Optical Unit Star Tracker DPU 5.4.4 GN&C Trades 5.4.5 GN&C Analytical Methods Basic Mass Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 142.61 17.4% 24.75 167.36 142.61 17.4% 24.75 167.36 7.10 135.51 5.0% 0.36 18.0% 24.39 7.46 159.90 608.77 18.0% 109.47 718.24 54.23 18.0% 9.76 63.99 54.23 18.0% 9.76 63.99 54.23 18.0% 9.76 63.99 229.25 9.4% 21.44 250.69 3.44 3.0% 0.10 3.54 3.44 3.0% 0.10 3.54 8 0.04 0.29 3.0% 0.01 0.30 2 0.58 1.16 3.0% 0.03 1.20 2 0.99 1.99 3.0% 0.06 2.05 The EDL profile was largely based on that of the Pioneer Venus large probe. The nominal profile can be seen in Figure 5.6. Accelerometers in the IMU are used to know when to trigger the deployment of the parachute. At a sufficiently low speed, roughly at a Mach of 0.7 and nominally just under 3 min from atmosphere entry, the heat shield has served its purpose and hence is jettisoned. A few seconds prior to heat shield jettison a parachute is deployed to create a sufficient difference in drag acceleration between the vehicle and the heat shield. A short time after the heat shield is jettisoned, at a time TBD, the landing legs are deployed. The time between heat shield jettison and landing leg deployment will probably be on the order of seconds to tens of seconds, basically just enough time to ensure that the heat shield has cleared the vehicle. After the landing legs have been deployed, at approximately 20 min after atmosphere entry, the parachute is released, which also releases the vehicle from the back shell. This is done to reduce the amount of drag on the vehicle and hence reduce the amount of time it takes for the vehicle to reach the surface. The vehicle then free falls for approximately another 70 min, reaching the surface roughly 90 min after atmosphere entry. To ensure that the landing load is less than the 40 g limit, the vehicle contains a drag flap to ensure a relatively low terminal velocity along with crush pads on the landing legs to absorb energy at impact. The Mission Analysis and Simulation Tool In Fortran (MASTIF) was used to simulate the nominal EDL profile for ALIVE. MASTIF contains a Venus atmosphere model, Venus-GRAM 2005, and was used to determine the required flight path angle that would provide a load no greater than 40 g’s. It was found that with the given assumptions at entry (altitude of 200 km, velocity of 11.3 km/s), a flight path angle of –8.7° was required to ensure that the maximum load experienced by the vehicle during atmospheric deceleration was less than 40 g’s. The nominal acceleration and altitude profile can be seen in Figure 5.7 and Figure 5.8, respectively. If the entry velocity can be reduced than the allowable flight path angle could be increased. 38 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 5.6—Summary of nominal EDL profile. Figure 5.7—Acceleration timeline from atmospheric entry. CD–2012-72 39 March 2012 COMPASS Final Report Figure 5.8—Nominal altitude profile during atmospheric entry. Parachute Sizing The goal when sizing the parachute was to create a drag area (Cd * Area) large enough that would cause a difference in drag acceleration on the vehicle and the heat shield such that no recontact would occur between the vehicle and heat shield after the heat shield was jettisoned. Table 5.8 shows the assumptions made during the parachute sizing process. It was felt that a difference in acceleration of about 4 m/s2 between the vehicle and the heat shield would be sufficient to ensure no re-contact after the heat shield was released. Table 5.8—Assumptions Made During Parachute Sizing Drag coefficients Heat shield ......................................................................................... 1.2 Vehicle (no chute, no heat shield) ...................................................... 1.0 Parachute ........................................................................................... 0.7 Parameters at time of chute deployment Vehicle velocity ........................................................................... 179 m/s Altitude .......................................................................................... 65 km 3 Atmospheric ......................................................................... 0.192 kg/m Parachute diameter Constructed diameter/inflated diameter ............................................. π/2 Sizing 2 Mass/constructed area ........................................................... 0.33 kg/m The drag force acting on the heat shield and the vehicle was calculated from the following equation: Drag Force = 0.5 ρv 2CdA where: ρ = atmospheric density v = air relative velocity Cd = Drag Coefficient A = projected area 40 Advanced Lithium Ion Venus Explorer (ALIVE) With the assumptions in Table 5.8, and assuming a 3.4 m diameter heat shield, the resulting drag on the heat shield after separating from the vehicle is 33.5 kN. Given that the mass of the heat shield is 371 kg, this results in a drag acceleration acting on the heat shield of 90.3 m/s2. This means that the acceleration on the vehicle with the inflated parachute, without the heat shield needed to be ~ 94.3 m/s2. Since the mass of the vehicle without the heat shield at the time of jettison is 1825 kg, this results in a required drag force of 173 kN. The required drag area (Cd*A) to produce 173 kN of force on the vehicle was then calculated to be 56.2 m2. The vehicle alone, without the parachute, contributes 9.1 m2 to the required drag area (Cd*A). Subtracting this 9.1 m2 of drag area from the required 56.2 m2 of drag area leaves 47.2 m2 left to be made up by the parachute itself. Assuming a drag coefficient of 0.7 for the parachute, this means that the required inflated area of the parachute is 67.4 m2, corresponding to an inflated diameter of 9.3 m. An assumption was then made that the inflated diameter would be a factor of π/2 smaller than the flat, constructed diameter. This resulted in a required constructed diameter of the parachute to be 14.5 m, corresponding to a total area of 166 m2. Once the cross sectional area was determined, the mass of the parachute was obtained by scaling the mass of the parachute used by the Galileo S/C since the subsystem lead had knowledge of both the cross sectional area and mass of that parachute. The ratio of mass to cross sectional area of the parachute used by the Galileo S/C was 0.33 kg/m2. With this knowledge, the mass of the parachute for ALIVE was then estimated to be 54 kg. 5.4.6 GN&C Risk Inputs At such a shallow flight path angle of –8.7°, there is an increased risk that the vehicle will not get captured by the atmosphere at the time of entry. Increasing the g-load limit would allow for a steeper flight path angle at entry, as would a lower entry velocity. Since the 11.3 km/s entry velocity was just an assumption at the time of this design, it is left as future work to iterate with the mission design lead to design an end to end trajectory that arrives at Venus with a lower entry velocity. 5.4.7 GN&C Recommendation As previously mentioned, no particular landing site was targeted by the GN&C subsystem. Atmospheric entry conditions were found that did in fact meet the 40 g load requirement for the EDL profile. It remains as future work however to iterate with the mission design lead to develop an end to end trajectory (interplanetary and EDL) that can deliver the vehicle to a specific, targeted landing site while meeting the less than 40 g load requirement. 5.5 Electrical Power System 5.5.1 Power Requirements Table 5.9 shows the power requirements for the specified mission stages. Ground operations and Launch, Cruise and Flyby and power needs are met by the solar arrays and Li-ion batteries. The Li-ion batteries are used for the Aeroshell/Parachute Descent and contained within a chamber that isn’t cooled but maintains acceptable temperatures during its multi-hour descent and duplex startup. Landed Science alternates between a “science” mode that operates for 6 hr continuous and requires 180 W of electrical power and “communications” mode that requires 380 W of power for 18 hr continuous. Because of this power fluctuation we use a combination of Li burner/Stirling and NaS batteries for power leveling. This allows us to operating the Stirling duplex at a constant electrical and cooling output while being able to follow the electrical power transients. Average electrical power is 330 W. CD–2012-72 41 March 2012 COMPASS Final Report Table 5.9—Power Requirements ALIVE total ALIVE total with 30% margin (W) Power (W) Lander with 30% margin Cruise deck with 30% margin Total lander and cruise deck with 30% margin 5.5.2 Ground ops and launch Deploy, cruise and flyby 8 hr 47.8 62.1 6480 hr 347.1 451.2 39.0 23.1 62.1 400.0 51.2 451.2 Drop cruise deck and Parachute aeroshell descent descent 1 hr 2 hr 301.7 384.9 392.2 500.4 392.2 N/A 392.2 500.4 N/A 500.4 Landed science mode Landed comm. mode 30 hr 138.7 180.3 90 hr 292.5 380.3 180.3 N/A 180.3 380.3 N/A 380.3 Power Assumptions The following assumptions were defined by the electrical power system lead for the ALIVE mission. Cruise Deck § § Body mounted arrays are practical for S/C cruise deck Li-ion batteries are located in a thermally isolated chamber without the need for separate cooling system along with a phase-change material to control temperature during descent. Lander § § § 5.5.3 Stirling Duplex can be integrated into Cold Box Li burner can transport its heat to Stirling while only losing 5% of its heat to surroundings A Stirling Duplex machine can be made which operates at heat to PV power efficiency of 50% of Carnot at a TR of 1.5. Power Design and MEL ALIVE Power System Design The ALIVE power system consists of two distinct parts. The first is the Cruise Deck power system and the second is the lander power system. The Cruise Deck uses body mounted solar arrays to provide power until the decent at Venus. Although Li-ion batteries are used for load leveling during the trip to Venus, these batteries are located on the Lander and used for descent power. The Lander power system has two distinct systems. The Li-ion batteries (also used during Cruise) power the vehicle during descent. Once on the surface a combination power and cooling by a Stirling duplex power that is driven by heat from the burning of Li and the Venus CO2 atmosphere. The engine/cooler system is assumed to be a conventional “Duplex Stirling” configuration in that the cooler and engine share the same mean operating pressure and frequency. Figure 5.9 shows a schematic of a Stirling Duplex. The convertor employs a simple monolithic heater head / pressure vessel. Figure 5.10 shows an overview of the heat and electrical flows of this single stage duplex system. The convertor hot end materials (MarM-243) are based upon those used in the Advanced Radioisotope Stirling Convertor (ASRG) with an upper temperature limit of 850 °C. The ASRG is currently creep life limited at 850 °C at 17 yr and for the short duration of this mission (5 d) we are projecting that an additional 100 °C (950 °C) will be our convertor upper temperature. The Li heat source is connected to the Stirling convertor using a sodium heat pipe. The Li burner is used to heat the gas inside the Stirling convertor that produces P-V work. Some of this work is converted to electrical power via a linear alternator while some of the P-V work drives the cooling stages. The advantage of the duplex system over separate Stirling power and cooling systems is rather than converting all of the PV power to electricity in the Stirling generator and then some 42 Advanced Lithium Ion Venus Explorer (ALIVE) back into piston motion for the cooler, we can use the work directly in the cooler (eliminating the alternator efficiency). Figure 5.9—Duplex Sketch Figure 5.10—Heat and power flows for a Stirling Duplex. Figure 5.11 shows the design point heat/power flows for the ALIVE Stirling duplex integrated with the cold box that contains the temperature sensitive electronics. This sketch shows an outer shell exposed to the ambient conditions and an inner shell containing the electronics and linear alternator. Approximately 14 kW of heat are generated by the burning of Li with the CO2 atmosphere. The products of this reaction are lower in density then the reactants and thus create a lower pressure area inside the tank drawing them into the Li tank. The insulation around the burner is sized to allow a 5% heat loss (665 W). Heat is transported to the Stirling duplex via a sodium heat pipe with the condenser being integrated into the Stirling duplex heater head. Approximately 13.3 kW of thermal power are put into the Stirling duplex to drive the cycle. Because of the low temperature ratio (TR) of the cycle (TR = 1.5, Thot = 950 °C, Tcold = 500 °C) Stirling convertors fraction of Carnot efficiencies are lower than that seen in other higher temperature ratio convertors (ASRG , TR>3). While ASRG has a fraction of Carnot efficiency approaching 60% it was assumed that this lower TR convertor would have a fraction of Carnot efficiency of only 50%. Overall PV efficiency was relatively low at 16%. Heat is rejected from the cycle via a pumped NaK loop. An electromagnetic pump (EMP) is used to move the liquid NaK over the cold end of the convertor and removing both the heat from the power generation portion of the system but also the heat from the cold box. Radiator area is 4.4 m2 and set by an assumed ∆T across the cold end of the CD–2012-72 43 March 2012 COMPASS Final Report convertor of 25 °C (maximum ∆T in order that cycle efficiency maximized) and 50 °C above the ambient environment (500 C). The EMP is 5% efficient. Electrical generation efficiency (after alternator and controller) was 15%. Average electrical power required by the system is 330 W. The majority of power generated in the duplex is used for cooling. Approximately 1500 W of PV power go into the cooler portion of the duplex. The cooler is assumed to be 35% of Carnot based on previous analysis of duplex cycles for the Venus atmosphere (VFDRM). Because the temperatures on the Venus surface are well above the allowable temperature of conventional magnets the linear alternator is placed within the cold box generating approximately 23 W of heat. Additionally, both heat led in from the environment and the heat generated from the electronics used to run the lander must also be removed. However, because communication power consumes a significant amount of power, much of the electrical power generated is emitted from the transmitters. Of the 330 W of electrical power generated only 165 W are added to the cold chamber with the rest emitted or used to charge the external load leveling batteries. High temperature NaS batteries are located on the external surface of the lander. After arrival at Venus, the NaS liquefies and the batteries start in a full state of charge. These batteries are used to start the duplex power system and take over for the Li-ion batteries after descent and landing. Figure 5.11—Heat and power flows for ALIVE Power and Cooling System. Figure 5.12 shows an overview of the lander along with the Duplex, Li tank and burner. Additionally the surrounding disk is a conceptual design of the pumped loop radiator that also serves as an additional drag to slow the descending S/C. 44 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 5.12—ALIVE Power/Cooling System highlights. Table 5.10 Mass Breakdown of Duplex Power System Mass (kg) Stirling Duplex ..................................................................................... 16 Burner and Insulation ............................................................................ 1 Radiator ............................................................................................... 22 Duplex Controller and PMAD ................................................................ 8 Li Tank................................................................................................ 5.5 Li Fuel (5 d) ....................................................................................... 200 EM Pump (pumped loop radiator) ......................................................... 2 Power Leveling Battery....................................................................... 4 Totals ................................................................................................ 259 Component All of the components of the power subsystem and their masses are shown in Table 5.11. Table 5.11—Electrical Power System ALIVE S/C MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.5 (kg) Chemical Power System Stirling Duplex Radiator Lithium Fuel and Tank PMAD Power Leveling Battery EM Pump Power Management & Distribution Burner Power Cable and Harness Subsystem (C and HS) Spacecraft Bus Harness 06.1.5.a.a 06.1.5.a.b 06.1.5.a.c 06.1.5.a.d 06.1.5.a.e 06.1.5.a.f 06.1.5.b 06.1.5.b.d 06.1.5.d 06.1.5.d.a 06.3.5 Unit Mass Lander Electrical Power Subsystem 06.1.5.a 06.3 QTY 06.3.5.c.a 06.3.5.c.b 06.3.5.d 06.3.5.d.a CD–2012-72 Total Mass (kg) (%) (kg) (kg) 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 277.50 12.2% 33.77 311.27 265.00 11.8% 31.27 296.27 16.00 16.00 20.0% 3.20 1 22.00 22.00 20.0% 4.40 1 213.30 213.30 10.0% 21.33 1 7.80 7.80 20.0% 1.56 1 3.90 3.90 20.0% 0.78 4.68 1 2.00 2.00 0.0% 0.00 2.00 1 0.50 0.50 1 12.00 12.00 1 25.00 1 3.00 12.00 1 5.00 20.0% 20.0% 20.0% 20.0% 0.10 19.20 26.40 234.63 9.36 0.60 0.10 2.40 0.60 14.40 2.40 14.40 229.25 9.4% 21.44 250.69 33.00 3.0% 1.00 34.00 28.00 0.0% 0.00 25.00 3.00 5.00 45 Growth 1917.94 0.50 Solar Array Power System Body Mounted Solar Array Batteries Power Cable and Harness Subsystem (C and HS) Spacecraft Bus Harness Growth 1 Cruise Deck Electrical Power Subsystem 06.3.5.c Basic Mass 5.00 0.0% 0.0% 20.0% 20.0% 28.00 0.00 25.00 0.00 3.00 1.00 6.00 1.00 6.00 March 2012 COMPASS Final Report Technology Maturity § § § 5.5.4 Solar arrays = TRL-6 Stirling Duplex = TRL-3 Li/CO2 Burner = TRL-3 Power Trades Power trades were performed on mission duration. Mission duration was varied until the landed Li and tank mass allowed the lander to fit within its mass limits. 5.5.5 Power Analytical Methods A spreadsheet Stirling duplex sizing tool that was developed for the radioisotope Venus duplex was used for this mission study. It was modified to add the burner and Li fuel and tank. 5.5.6 Power Risk Inputs The following are the power risks: § § § § 5.5.7 Unable to make a Stirling power portion operate as 50% of Carnot at a temperature ratio of 1.5. Unable to make a Stirling cooler operate at 35% of Carnot. Unable to effectively integrate Li burner/Stirling duplex Unable to create a closed (i.e., no release to atmosphere) Li/ CO2 burner Power Recommendation The following are the future work and recommendations from the power subsystem lead: § § § 5.6 More detailed design of the heat pipe to Stirling duplex interface Preliminary design of Stirling duplex to ensure regenerator length can match insulation thickness requirements. Consider higher temperature electronics Propulsion System (Entire section 5.6 and subsections provided by the Power Seat, except for the MEL, provided by the Systems Integration Lead) 5.6.1 Propulsion System Requirements The propulsion system is required to provide adequate total impulse, at an acceptable thrust level, to perform trajectory adjustments and maintain proper vehicle orientation during the cruise to Venus. Prior to jettisoning the cruise stage, the propulsion system is required to orient the S/C to the desired orientation for Venus atmospheric entry. In order to reduce risk and cost, the propulsion system is required to be single fault tolerant, and composed of high TRL level COTS components. Finally, propellant is to be stored and provided at the conditions and flow rates required by the propulsion system, regardless of the number of thrusters firing at any given time. 5.6.2 Propulsion System Assumptions It is assumed that a single fault tolerant hydrazine based blow down system is used. It is also assumed that small thrusters are used for S/C orientation, while larger thrusters in the axial direction are used for trajectory adjustments. 46 Advanced Lithium Ion Venus Explorer (ALIVE) 5.6.3 Propulsion System Design and MEL The entire propulsion system is located on the cruse deck, which is jettisoned prior to Venus atmospheric entry. The system is comprised of sixteen thrusters, located in four clusters containing four thrusters each, two nitrogen pressurized commercial off-the-shelf membrane tanks, and a single fault tolerant feed system. Each cluster of thrusters contains three MR-103C thrusters which can deliver 0.9 N (0.2 lbf) of thrust at a nominal ISP of 220 s, and are used to provide fine attitude control. Each cluster also has one larger MR106E thruster delivering 22.3 N (5.0 lbf) of thrust at a nominal Isp of 230 s, and are used to provide axial thrust. All four clusters are feed hydrazine propellant via a single fault tolerant feed system comprised of various COTS components, a nominal instrumentation suite including Pain Electronics flight certified pressure sensors and thermocouples, tank and line heaters, and MLI. The system is fueled via a set of Vacco V1E10430-01fill and drain valves, which are flight qualified and have a metal to metal primary seat. The propellant is filtered via Vacco F1D10638-01 15 µm absolute propellant filters. Tank isolation is provided by three MOOG 51-166 valves, although pyrotechnic valves could be substituted. The hydrazine is stored in two ATK 80275-1 Ti alloy (Ti-6Al-4V) spherical membrane tanks with a volume of 37.69 L (2300 in3) and a MOP of 30 bar (435 psia). Some of the feed system components are shown in Figure 5.13, and a preliminary P&ID of the system is shown in Figure 5.14. Figure 5.13—Feed System components. Figure 5.14—Preliminary Cruse Deck Propulsion P&ID CD–2012-72 47 March 2012 COMPASS Final Report The total propellant mass is calculated using information from both the trajectory mission analysis output, as well as internal propellant and propulsion system calculations. The three different propellants tracked in the MEL are: Used, Residuals, and Performance Margin. These are defined as follows: Used.—The used propellant is calculated using an ideal equation. This is the propellant necessary to push the mass of the S/C using the total mission ΔV and the idealized form of the rocket equation. There is no margin on the used propellant. Performance Margin.—The performance margin is calculated by taking a percentage of the propellant use for total ΔV performed by that particular propulsion system. For this analysis, 10% is used. Residuals.—The residuals are calculated by taking the total mass of the used and margin propellants, and calculating a percentage of that mass. For this analysis, 3.5% is used to calculate the residual hydrazine mass. Total propellant.—The total propellant of the mission is the sum of used, margin and residuals. mTotalPropellant = mUsed + mMargin + mResiduals These divisions of propellant are used in the calculation of dry, wet and inert mass of the total S/C. A listing of all major propulsion system component masses as captured in the MEL shown in Propulsion System Trades There were no propulsion system trades conducted for this study. 5.6.4 Propulsion System Analytical Methods The methods used to design the propulsion system involve using a mix of published values, empirical data, and analytical tools. Published values and empirical data are used wherever possible, with analytical tools being used as necessary. These include National Institute of Standards and Technology (NIST) tables, CEA, and other fluid/gas property codes, as well as custom tools developed form basic physical relationships and conservation equations with empirical based inclusions for real life hardware requirements (mounting bosses, flanges, etc.). Table 5.12. 5.6.5 Propulsion System Trades There were no propulsion system trades conducted for this study. 5.6.6 Propulsion System Analytical Methods The methods used to design the propulsion system involve using a mix of published values, empirical data, and analytical tools. Published values and empirical data are used wherever possible, with analytical tools being used as necessary. These include National Institute of Standards and Technology (NIST) tables, CEA, and other fluid/gas property codes, as well as custom tools developed form basic physical relationships and conservation equations with empirical based inclusions for real life hardware requirements (mounting bosses, flanges, etc.). Table 5.12—Propulsion System ALIVE S/C MEL 48 Advanced Lithium Ion Venus Explorer (ALIVE) WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design QTY Unit Mass (kg) Cruise Deck Propulsion (Chemical Hardware) 06.3 06.3.7 06.3.7.a 06.3.7.a.b 06.3.7.a.b.b 06.3.7.a.b.c 06.3.7.a.b.d 06.3.7.b 06.3.7.b.b 06.3.7.b.b.a 06.3.7.b.b.f Primary Chemical System Hardware Reaction Control System Hardware RCS Thruster Subassembly Large RCS Thrusters Small RCS Thrusters Propellant Management (Chemical) RCS Propellant Management Fuel Tanks Feed System - regulators, valves, etc Basic Mass Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 229.25 9.4% 21.44 250.69 30.52 5.2% 1.58 32.10 11.04 4.9% 0.54 11.58 11.04 4.9% 0.54 11.58 4 0.50 2.00 18.0% 0.36 2.36 4 1.27 5.08 2.0% 0.10 5.18 12 0.33 3.96 2.0% 0.08 19.48 5.3% 1.04 19.48 5.3% 1.04 2 7.71 15.42 1 4.06 4.06 4.04 20.52 20.52 2.0% 0.31 15.73 18.0% 0.73 4.80 Thrust requirements and propellant load are determined by GN&C analysis. Using those results, the tanks are selected so that both adequate propellant and tank pressure are available to ensure proper propulsion system performance during the entire mission, and that adequate engine performance is available to meet both vehicle and mission requirements and constraints. 5.6.7 Propulsion System Risk Inputs One constant risk with hydrazine is the possibility of it freezing, especially on the shadow side of the S/C, which could cause a loss of mission. Detailed thermal analysis, however, can provide MLI and strip heater power requirements that minimize this risk. 5.6.8 Propulsion System Recommendation Since the propellant tanks are COTS, they are slightly oversized for their respective propellant loads. Therefore, it is recommended that the hydrazine tanks be filled to capacity to provide additional delta-V margin, assuming that this doesn’t negatively impact S/C wet mass and/or LV launch margin to an unacceptable degree. Another recommendation is to conduct a propellant trade of hydroxyl-ammonium nitrate (HAN) based monopropellants versus hydrazine. Although this mission doesn’t really require the cold temperature capability of the HAN monopropellants, their lack of toxicity relative to hydrazine may lower ground handling related costs. As of this writing, however, HANs are still undergoing materials compatibility testing, and thus may be too risky for this class of mission in the near term. 5.7 Structures and Mechanisms (Entire section 5.8? and subsections provided by the Structures Seat, except the MEL, which is provided by Systems Integration Lead) 5.7.1 Structures and Mechanisms Requirements The S/C must contain the necessary hardware for research instrumentation, avionics, communications, power, and propulsion. It must be able to withstand applied loads from the LV, landing on the Venus surface, and operating in the Venetian environment. The maximum axial acceleration of 44 g (430 m/s2, 1420 ft/s2) is during descent to the Venetian surface. The Venus surface is at approximately 480 °C (900 °F) in temperature and 9 MPa (1300 psi) pressure. In addition, the S/C bus has to provide minimum deflections, sufficient stiffness, and vibration damping. Weight has to be kept to a minimum and the stowed S/C must fit the confines of the LV. Mechanisms are used to separate from the LV, jettison the heat shield, deploy landing legs, and jettison the backshell. CD–2012-72 49 March 2012 COMPASS Final Report 5.7.2 Structures and Mechanisms Assumptions The S/C bus provides the main backbone for the S/C. It is constructed of a Ti alloy, Ti-6Al-4V. The Cruise Deck is a simple frustum, also, constructed of the Ti alloy, Ti-6Al-4V. The Ti alloy, used in the construction of the S/C, is specified in the Federal Aviation Administration’s Metallic Materials Properties Development and Standardization (MMPDS) (2006). The main bus consists of a sphere and strut mounted hardware. 5.7.3 Structures and Mechanisms Design and MEL The main bus of the Lander consists of a sphere, which provides the most efficient approach for surviving the Venus environment while keeping mass to a minimum. Secondary components, such as struts and mounting flanges/rings consist of Ti also. The fuel container is cylindrical. The inside of the container is exposed to the Venetian atmospheric pressure. This negates the need for thick walls relative to the main spherical bus. A ring flange, mounted to the top of the tank, is utilized to attach the support struts from the S/C to the tank. A smaller Ti sphere is used to house the science instruments. A mounting ring is located equatorially around the science sphere and is used to attach the struts that support the sphere to the S/C. Landing gear consists of rigid tubular members. The main tube of each landing leg has a lockable hinge to allow stowing the landing gear within the aeroshell assembly. Crushable Ti honeycomb, mounted to the base of each pad, is used to absorb the energy upon landing on the surface of Venus. The honeycomb is a commercial component, Benecor, Inc. Ti3AL2.5V Honeycomb 9.56 (.125/.002). Tubular members support and attach the radiators to the S/C. Similarly, ring flanges, ribs, and tubular struts are used to mount aero drag flaps to the S/C. Figure 5.15 illustrates the Lander in stowed and deployed states. Pyrotechnic fasteners are specified for all the separation planes. The devices provide a simple, reliable, and light weight approach for handling the separation of the various components. Table 5.13 shows the expanded MEL for the structures subsystem on the EZE Lander platform. This MEL breaks down the structures line elements to the lowest WBS. (a) (b) Figure 5.15—(a) The Lander stowed within the heat shield/backshell assembly and (b) the Lander fully deployed. 50 Advanced Lithium Ion Venus Explorer (ALIVE) Table 5.13—ALIVE S/C Structures MEL QTY WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design Unit Mass (kg) Lander Structures and Mechanisms 06.1 06.1.11 Structures Mechanisms 06.1.11.a 06.1.11.b Aeroshell Structures and Mechanisms 06.2 06.2.11 Structures Mechanisms 06.2.11.a 06.2.11.b Cruise Deck Structures and Mechanisms 06.3 06.3.11 Structures Mechanisms 06.3.11.a 06.3.11.b Basic Mass Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 513.91 18.0% 92.50 606.42 491.23 18.0% 88.42 579.65 22.68 18.0% 4.08 26.76 608.77 18.0% 109.47 718.24 181.85 18.0% 32.73 214.58 150.44 18.0% 27.08 177.51 31.41 18.0% 5.65 37.07 229.25 9.4% 21.44 250.69 88.01 18.0% 15.84 103.86 75.88 18.0% 13.66 89.54 12.13 18.0% 2.18 14.31 Basic Mass Growth Growth Total Mass Table 5.14—Lander Structures MEL QTY WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.11 Unit Mass (kg) Lander Structures and Mechanisms 06.1.11.a 06.1.11.a.a Structures Primary Structures 06.1.11.a.a.a Primary structure, sphere 06.1.11.a.a.b 06.1.11.a.a.c 06.1.11.a.a.d 06.1.11.a.b 1 234.95 Flange assy., sphere middle 1 22.50 Ring, hardware mounting 1 8.34 Sphere, science 1 56.96 Secondary Structures (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 1079.92 16.4% 177.57 1257.49 513.91 18.0% 92.50 606.42 491.23 18.0% 88.42 579.65 322.75 18.0% 58.10 234.95 18.0% 380.85 42.29 22.50 18.0% 4.05 8.34 18.0% 1.50 56.96 18.0% 10.25 168.48 18.0% 30.33 277.24 26.55 9.84 67.22 198.80 06.1.11.a.b.a Fuel tank mount assembly 1 19.63 06.1.11.a.b.b Landing gear assembly 1 135.83 06.1.11.a.b.c Radiator support 1 1.06 1.06 18.0% 0.19 1.26 06.1.11.a.b.d Science sphere mounts 1 2.62 2.62 18.0% 0.47 3.09 Flange, heat shield to fuel tank 1 9.34 1.68 11.03 06.1.11.a.b.e 06.1.11.b 06.1.11.b.f Mechanisms Installations 19.63 18.0% 3.53 135.83 18.0% 24.45 9.34 18.0% 23.16 160.28 22.68 18.0% 4.08 26.76 22.68 18.00% 4.08 26.76 06.1.11.b.f.b ECLSS Installation 1 1.59 1.59 18.00% 0.29 1.88 06.1.11.b.f.c GN&C Installation 1 5.70 5.70 18.00% 1.03 6.73 1.07 06.1.11.b.f.d Command and Data Handling Installation 1 0.90 0.90 18.00% 0.16 06.1.11.b.f.e Communications and Tracking Installation 1 1.95 1.95 18.00% 0.35 2.30 06.1.11.b.f.f Electrical Power Installation 1 11.10 11.10 18.00% 2.00 13.10 06.1.11.b.f.g Therrmal Control Installation 1 1.43 1.43 18.00% 0.26 1.69 Table 5.15—Aeroshell Structures MEL CD–2012-72 51 March 2012 COMPASS Final Report QTY WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design Basic Mass Unit Mass (kg) 06.2.11 06.2.11.a.a 06.2.11.a.a.a Aeroshell back 1 (%) (kg) (kg) 16.1% 308.47 2226.41 135.86 608.77 18.0% 109.47 718.24 181.85 18.0% 32.73 214.58 150.44 18.0% 27.08 177.51 135.86 18.0% 24.46 160.32 135.86 Secondary Structures 06.2.11.a.b Total Mass (kg) Structures Primary Structures 06.2.11.a Growth 1917.94 Aeroshell Structures and Mechanisms 06.2 Growth 14.57 18.0% 18.0% 24.46 2.62 160.32 17.19 06.2.11.a.b.a Flange, aeroshell back to chute housing 1 5.23 5.23 18.0% 0.94 6.17 06.2.11.a.b.b Flange, heat shield to fuel tank 1 9.34 9.34 18.0% 1.68 11.03 Mechanisms Adaptors and Separation 06.2.11.b 06.2.11.b.e 31.41 18.0% 5.65 37.07 14.40 18.00% 2.59 16.99 06.2.11.b.e.a Pyrotechnic fasteners & springs, heat shield 6 1.20 7.20 18.00% 1.30 06.2.11.b.e.c Pyrotechnic fasteners & springs, back shell 6 1.20 7.20 18.00% 1.30 Installations 06.2.11.b.f 17.01 06.2.11.b.f.c GN&C Installation 1 2.16 06.2.11.b.f.g Therrmal Control Installation 1 14.85 18.00% 3.06 8.50 8.50 20.07 2.16 18.00% 0.39 2.55 14.85 18.00% 2.67 17.53 Table 5.16—Cruise Deck Structures MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design QTY Unit Mass (kg) Cruise Deck Structures and Mechanisms 06.3 06.3.11 06.3.11.a 06.3.11.a.a 06.3.11.a.a.a 06.3.11.a.b 06.3.11.a.b.a 06.3.11.b 06.3.11.b.e 06.3.11.b.e.a Structures Primary Structures Main Cruise Deck Structure 1 70.66 1 5.23 Secondary Structures Mechanisms Adaptors and Separation Pyrotechnic fasteners & springs 6 1.20 Growth Growth Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 229.25 9.4% 21.44 250.69 88.01 18.0% 15.84 103.86 75.88 18.0% 13.66 89.54 70.66 18.0% 12.72 70.66 5.23 Flange, aeroshell back to chute housing Installations 06.3.11.b.f Basic Mass 5.23 18.0% 18.0% 18.0% 12.72 0.94 0.94 83.38 83.38 6.17 6.17 12.13 18.0% 2.18 14.31 7.20 18.00% 1.30 8.50 7.20 4.93 18.00% 18.00% 1.30 0.89 8.50 5.82 06.3.11.b.f.c GN&C Installation 1 0.14 0.14 18.00% 0.02 0.16 06.3.11.b.f.f Electrical Power Installation 1 1.32 1.32 18.00% 0.24 1.56 06.3.11.b.f.i Chemical Propulsion Installation 1 3.47 3.47 18.00% 0.62 4.10 5.7.4 Structures and Mechanisms Trades No trades for structural design were considered for this study. 5.7.5 Structures and Mechanisms Analytical Methods The high pressure and temperature of the atmosphere on the surface of Venus provides challenges for maintaining the structural integrity of a Lander. All the main structural components are fabricated from the Ti alloy, Ti-6Al-4V. The high pressure environment causes potential issues with buckling of the structure. The sphere of the main bus was checked for buckling and the wall thickness was specified to minimize the risk. The equation, presented by Young’s and Budynas’ Roark’s Formulas for Stress and Strain (2002), for determining the external pressure for buckling a sphere is (1) The equation represents a probable actual minimum pressure to cause buckling. The variables from the equation are 52 Advanced Lithium Ion Venus Explorer (ALIVE) P = pressure to cause buckling E = Young’s modulus of the material t = wall thickness of the sphere r = radius of the sphere Solving Equation (1) for the wall thickness and applying a safety factor of 1.5 results in a minimum wall thickness of 12 mm (0.47 in). The sphere for the science instruments has the same wall thickness as the main bus sphere. The original drag flap design had the supports cantilevered out from the center. The expected 2000 kg mass at the given stage of the trajectory and 44 g (430 m/s², 1411 ft/s²) deceleration significantly exceeded the strength limits of the structure. As a result, support struts were added around the outer perimeter of the drag flaps. The crushable honeycomb pads on each leg of the landing gear were sized to limit the deceleration to 40 g (390 m/s2, 1280 ft/s2) upon landing. The approach velocity is estimated to be 6.3 m/s (250 in/s). Using the physics equations of motion the resulting necessary displacement of the crushable honeycomb pads is a minimum of 0.051 m (2.0 in). Assuming the landing load is distributed evenly among the three landing legs the force per leg is 141 kN (31,700 lbf). The necessary diameter of each pad is 312 mm (12.3 in) for a Ti honeycomb that has a high temperature ultimate strength of 5.76 MPa (835 psi). The honeycomb pads are sized to have the applied load induce a stress at the approximate ultimate strength of the honeycomb. A quick check was made to size the lower standoffs between the spheres of the double walled main bus structure. The inner sphere and its contained hardware were estimated to be 100 kg (220 lb). A maximum of 200 g (1960 m/s2, 6430 ft/s2) is anticipated. Four supports or standoffs at 30° from the vertical are assumed for the lower support. Using tubes of 5 cm (2.0 in.) OD with 3 mm (0.12 in.) thick walls the resulting maximum stress is approximately 128 MPa (18.5 ksi). The yield strength of Ti-6Al-4V is approximately 530 MPa (77 ksi) as per the Federal Aviation Administration’s MMPDS (2006). Using a safety factor of 1.5 provides a material limit of 350 MPa (51 ksi). The resulting margin is 1.7. An additional installation mass was added for each subsystem. These installations were modeled using 4% of the CBE dry mass of each of the subsystems. The 4% magnitude for an initial estimate compares well with values reported by Heineman (1994) for various systems. This is to account for attachments, bolts, screws and other mechanisms necessary to attach the subsystem elements to the bus structure and not book kept in the individual subsystems. 5.7.6 Structures and Mechanisms Risk Inputs Structural risks may include excessive g loads, impact from a foreign object, or harsh landing on Venus which may cause too much deformation, vibrations, or fracture of sections of the support structure. Consequences include lower performance from mounted hardware to loss of mission. Excessive deformation of the structure can misalign components dependent on precise positioning, therefore, diminishing their performance. Internal components may be damaged or severed from the rest of the system resulting in diminished performance or incapacitation of the system. Excessive vibrations may reduce instrumentation performance and/or potentially lead to long term structural failure due to fatigue. Overall, the mission may not be completed in an optimum manner or it can be terminated in the worst case. In an effort to mitigate the structural risk the structure is to be designed to NASA standards to withstand expected g loads, a given impact, and to have sufficient stiffness and damping to minimize issues with vibrations. Trajectories are to be planned to minimize the probability of impact with foreign objects. CD–2012-72 53 March 2012 COMPASS Final Report Similar to the structural risks excessive g loads, impact from a foreign object, or harsh landing may damage mechanisms. Consequences include lower performance from mounted hardware to loss of mission. Failure of mechanisms may prevent optimum hardware operation or may inhibit mission completion. Failure of separation or deployment units can prevent planned mission completion. Mitigation of the risks with mechanisms would include the mechanisms are to be designed to NASA standards to withstand expected environmental conditions. All precautions should be taken to prevent damage from installation, launch, and operating conditions. 5.7.7 Structures and Mechanisms Recommendation Mass savings may be realized with different materials and architectures. Although, the harsh environment presented by Venus may limit material selection. Sandwich construction composites, isogrids, or orthogrids may be considered. A detailed stress analysis using numerical methods may be applied to optimize the design for the anticipated mission loads. 5.8 Thermal Control The thermal control system for the Venus lander mission is broken down in the thermal control for the various segments of the mission, transit to Venus, entry into the Venus atmosphere and operation on the Venus surface. The thermal control system for each stage in the mission is described in the following sections. 5.8.1 Cruise Deck Thermal Control The cruise deck thermal control system has to protect and regulate the temperature of the S/C and lander as it transits from Earth to Venus. The Stirling cooler cools the components within the lander during transit. The heat removed by the cooler must be rejected to space through the use of a radiator on the cruise deck. The environment in which the thermal control system has to operate to maintain the desired internal operating temperature of the electronics and lander varies from near Earth operation to deep space transit to operation near Venus. The sizing of the components of the thermal system is based on operation within this environment. The heat transfer to and from the S/C is based on a radiative energy balance between the vehicle and its surroundings. Solar radiation is the main source of external heat for the majority of the mission, during transit. Operation near Earth and Venus also involves the albedo (reflected sunlight) from the planet as well as direct radiation (infrared (IR)) from the planet itself. These environmental conditions are listed in Table 5.17. Table 5.17—Transit Environment Constants Constant Solar Intensity Albedo Planet IR Earth 2 1360 W/m 0.3 2 240 W/m Venus 2 2613 W/m 0.75 2 141 W/m To maintain the S/C and lander components at their desired operating temperature the following components were utilized for the cruise deck thermal control. § § § § § Electric heaters, thermocouples and data acquisition for controlling the temperature of the electronics. MLI for insulating the electronics and temperature sensitive components. Thermal paint for minimal thermal control on exposed structural surfaces. Radiator for rejecting heat from the enclosed lander. Cold plates with heat pipe connections to the radiator, for channeling the heat from the lander to the radiator. 54 Advanced Lithium Ion Venus Explorer (ALIVE) 5.8.2 Electric Heaters The electric heaters were used to provide added thermal control to the cruise deck electronics during transit. Strip heaters, as shown in Figure 5.16, were used to provide heat to the reaction control system propellant lines and other components within the cruise deck. Thermal control is accomplished through the use of a network of thermocouples whose output is used to control the power to the various heaters. A data acquisition and control computer is used to operate the thermal system. Figure 5.16—DuPont Kapton Strip Heater. The mass breakdown of the thermal system for the ALIVE is shown in Table 5.18. Table 5.18—Thermal ALIVE S/C MEL WBS Description Number Case 1 NIAC Venus Spacecraft CD-2012-72 06 Extended Venus Explorer Spacecraft Design 06.1 06.1.6 Total Mass (kg) (%) (kg) (kg) 1917.94 16.1% 308.47 2226.41 16.4% 177.57 1257.49 6.44 42.23 Active Thermal Control Data Acquisition Thermocouples Passive Thermal Control Heat Sinks Heat Pipes Electronics Enclosure Insulation 1.50 18.0% 0.27 06.1.6.b.a 06.1.6.b.b 06.1.6.b.c 1 1.00 5 0.10 Passive Thermal Control Ablative Material 06.2.6.b.a 06.3.6.a.c 06.3.6.a.d 06.3.6.b 06.3.6.b.c 06.3.6.c 06.3.6.c.c 18.0% 18.0% 1.77 0.18 0.09 6.17 1.18 0.59 40.46 4 0.14 0.55 18.0% 0.10 0.65 0.21 0.84 18.0% 0.15 0.99 1 32.89 1 Active Thermal Control Thermal Controller Data Acquisition Thermocouples Passive Thermal Control Electronics Enclosure Insulation Semi-Passive Thermal Control (cruise deck and internal) Radiator 06.3.6.a.b 0.50 18.0% 4 371.29 Cruise Deck Thermal Control (Non-Propellant) 06.3.6.a 1.00 34.29 Aeroshell Thermal Control (Non-Propellant) 06.2.6.b 5.8.1 Growth 18.0% 06.1.6.b 06.3.6 Growth 35.79 06.1.6.a.d 06.3 (kg) Basic Mass 1079.92 06.1.6.a.c 06.2.6 Unit Mass Lander Thermal Control (Non-Propellant) 06.1.6.a 06.2 QTY 32.89 18.0% 608.77 18.0% 371.29 371.29 371.29 5.92 38.81 109.47 718.24 18.0% 66.83 438.13 18.0% 66.83 438.13 18.0% 229.25 9.4% 10.34 2.90 66.83 438.13 21.44 250.69 18.0% 1.86 12.20 18.0% 0.52 3.42 2 0.20 0.40 18.0% 0.07 0.47 2 1.00 2.00 18.0% 0.36 2.36 5 0.10 0.09 0.59 0.50 5.56 1 5.56 5.56 1.88 1 1.88 1.88 18.0% 18.0% 18.0% 18.0% 18.0% 1.00 6.56 1.00 0.34 6.56 2.21 0.34 2.21 MLI and Thermal Control Paint MLI was used to insulate the cruise deck electronic components and exposed propellant tands to minimize their heat loss for deep space operation. MLI is constructed of a number of layers of metalized material with a nonconductive spacer between the layers. The metalized material has a low absorptivity that resists radiative heat transfer between the layers. The insulation can be molded to conform over the exterior of the cruise deck or any individual component, as shown in Figure 5.17. CD–2012-72 55 March 2012 COMPASS Final Report Figure 5.17—Example of MLI blanket design and application. In exposed areas where MLI cannot be applied, mainly exposed structural components, thermal control paint is applied. Since the S/C will be exposed to direct sunlight for the majority of its operation, this paint is used to minimize the absorption of solar radiation. This helps maintain thermal control of the vehicle by minimizing the temperature of exposed components. The paint utilized is AZ-93. Its characteristics are listed in Table 5.19. Table 5.19—MLI Specifications Variable Value MLI Emissivity................................................................................... 0.07 MLI Material ................................................ Metalized (Al) Kapton layers Layer Thickness ...................................................................... 0.025 mm Number of MLI layers .......................................................................... 25 AZ-93 Emissivity ............................................................................... 0.91 AZ-93 Absorptivity ............................................................................ 0.15 5.8.2 Radiator and Cold Plates To reject heat from the lander during transit from the Earth to Venus, a radiator was utilized. This radiator was coupled to the hot end of the Stirling cooler through a cold plate interface. The Stirling cooler was used to remove any waste heat from the interior of the lander during transit. Heat pipes were used to move heat from the cold plate to the radiator panel, which then rejected the heat to space. An example of a cold plate with integral heat pipes is shown in Figure 5.18. The radiator was sized for operation near Venus. This is the worst case operating condition for rejecting heat from the radiator. The radiator was coated to limit its solar radiation absorption characteristics. The details on the radiator sizing are given in Table 5.20. The radiator was surface mounted to the cruise deck and therefore rejected heat from one side. The radiator was sized based on an energy balance approach, utilizing the thermal heat needed to be rejected and the incoming thermal radiation from Venus and the sun. An example of a S/C radiator with integral heat pipes is shown in Figure 5.19. 56 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 5.18—Example of a cold plate with integrated heat pipes. Table 5.20—Cruise Deck Radiator Sizing Component Value Radiator Solar Absorptivity .............................................................. 0.14 Radiator Emissivity ......................................................................... 0..84 Estimated Maximum Radiator Solar Angle ....................................... 70° Total Radiator Dissipated Thermal Power ................................... 152 W View Factor to Venus ....................................................................... 0.25 2 Required Radiator Area .............................................................. 0.24 m Radiator Operating Temperature ................................................... 358 K Cold Plate Material............................................................................... Al Cold Plate Dimensions........................................... 0.1- by 0.1- by 5-mm Figure 5.19—Radiator with integral heat pipes (ACT, inc). 5.9 Venus Atmospheric Environment The harsh environment of Venus provides a number of challenges in the operation of equipment and materials. Operating within this environment, from entry to descent to operation on the surface requires significant thermal control. The atmosphere is composed of mainly CO2 but does contain corrosive components such as sulfuric acid. The planet has a very thick atmosphere and is completely covered with clouds. The temperature and pressure near the surface is 455 °C at 90 Bar. The atmospheric properties (temperature, wind speed, solar attenuation and atmospheric density) from the surface to 100 km altitude are shown in Figure 5.20 and illustrated in Figure 5.21. CD–2012-72 57 March 2012 COMPASS Final Report 500 100 Temperature Wind Velocity Attenuation Density 400 90 300 70 60 200 50 100 40 30 0 0 10 20 30 40 50 60 70 80 90 100 20 -100 10 -200 0 Altitude (km) Figure 5.20—Venus atmospheric properties. Figure 5.21—Venus atmospheric structure. 58 Attenuation (%), Density (kg/m^3) 80 Advanced Lithium Ion Venus Explorer (ALIVE) The winds within the atmosphere blow fairly consistently in the same direction as the planetary rotation (East to West) over all latitudes and altitudes up to 100 km. Above 100 km, the winds shift to blow from the dayside of the planet to the night side. The wind speeds decrease as a function of altitude from ~100 m/s at the cloud tops (60 km) to ~0.5 m/s at the surface. These high wind speeds and the slow rotation of the planet produce a super rotation of the atmosphere (nearly 60 times faster than the surface). These high wind speeds and the slow rotation of the planet produce a super rotation of the atmosphere (nearly 60 times faster than the surface). 5.10 Aeroshell and Descent Thermal Control The aeroshell consists of a heat shield and back shell. The heat shield needs to be able to withstand the aerodynamic heating that will be encountered during entry into the Venus atmosphere. The heat is generated by friction caused by the drag of the capsule as it enters the atmosphere. The heat load will depend on the entry angle and speed. The heat shield for Venus entry was scaled off of the Stardust and Genesis Earth entry vehicles as well as the proposed Orion entry vehicle. All of these vehicles had similar entry velocities (~ 11 km/s) to what is expected for the Venus lander aeroshell. The heat shield sizing utilized the Orion structural design, shown in Figure 5.22, but substituted AVCOAT for PICA as the ablative material. This was done due to the size of the heat shield. The AVCOAT thickness utilized was 4.3 cm. The materials breakdown for the heat shield is given in Table 5.21. Figure 5.22—Orion heat shield structural makeup. Table 5.21—Heatshield Material Layer Properties Material Thickness (cm) 4.3 0.0305 0.229 0.0305 0.102 4.83 0.102 Avcoat RTV Glue Foam Insulation (SIP) RTV Glue Ti Alloy (Ti-6Al-4V) Ti Alloy (Ti-3Al-2.5V) Honeycomb Ti Alloy (Ti-6Al-4V) Density 3 (kg/m ) 510 1060 70 1060 4430 96.3 4430 The heat shield and backshell geometry were scaled up from the Stardust aeroshell design (shown in Figure 5.23). The Stardust aeroshell and entry specifications are: § § § § § § Entry velocity was 11.04 km/s 60° half angle –8.0° entry angle, 15 rpm 4 hr before entry Backshell thickness 5 cm Heat shield/structure thickness 10 cm CD–2012-72 59 March 2012 COMPASS Final Report Figure 5.23—Stardust Aeroshell Geometry 5.10.1 Descent Electronics Enclosure Thermal Control The descent electronics enclosure is an insulated pressure vessel that contains the electronics, equipment and sensors that are utilized during decent and landing. The enclosure does not have any active cooling. It utilizes aerogel insulation and phase change material to maintain the internal temperature of the enclosure at approximately 300 K during the descent for duration of 1 hr, as illustrated in Figure 5.24. To maintain the interior temperature of the enclosure, a layer of aerogel insulation is utilized on the inside of the pressure vessel outer wall. On the inside of the insulation is a layer of phase change material. It was selected because of its melting point of 305 K. As heat enters the chamber through the insulation it will cause the phase change material to melt. For the 1 hr descent all of the thermal energy leaking in through the insulation will be absorbed by the sodium sulfate through a phase change between a solid and liquid. This will maintain the interior temperature of the chamber at around 305 K. The specifications for the thermal control components for the descent electronics enclosure are given in Table 5.22. Figure 5.24—Descent electronics thermal control items. Table 5.22—Insulation and Phase Change Material Specifications Item Material Thickness Density Mass Insulation Aerogel 2 cm 3 20 kg/m 0.4 kg 60 Phase change material Sodium sulfate 7 mm 3 1464 kg/m 22.5 kg Advanced Lithium Ion Venus Explorer (ALIVE) 5.11 Surface Lander Thermal Control All of the components that require a low temperature, relative to the atmosphere, for operation are located within the electronics enclosure pressure vessel. This pressure vessel is actively cooled by the Stirling cooler system. To minimize the power needed to cool this enclosure it is insulated from the outside environment. Within the pressure vessel along the outer surface wall is aerogel insulation. This insulation is utilized to reduce the heat leak in from the external atmospheric conditions. The exterior temperature was assumed to be 735 K and the inside operational temperature was 300 K. In addition to heat leaking in through the insulation, heat also entered through a number of penetrations through the insulation that were necessary for the vehicle operation. These included wires, view ports and structural support standoffs. The heat leak into the chamber came from a number of sources. The interior of the pressure vessel was at 1 atm. Utilizing a gas within the pressure vessel provided a number of benefits. It allowed more even heat transfer between the electronics and the Stirling cooler. Also since the insulation selection and designed was made to operate within a 1 atm environment its operation was less susceptible to small leaks into the pressure vessel. If a completely evacuated pressure vessel was utilized along with MLI, any gas leak into the chamber would significantly reduce the insulation’s insulating capability and could be mission ending. However, with the aerogel insulation, it is capable of operating over a much larger pressure range and therefore is not very sensitive to minor leaks of gas into the pressure vessel. Also if atmospheric gas was to leak into the pressure vessel at a slow rate, there would be a slow degradation of the insulating capability of the aerogel which would mean a reduced mission time as the temperature slowly rose within the chamber but not a catastrophic mission failure as would occur in a similar situation with MLI. A diagram of the heat leak rates into the pressure vessel through the various components is shown in Figure 5.25 and the characteristics of each is given in Table 5.23. Figure 5.25—Heat leak into the Lander electronics enclosure pressure vessel. Table 5.23—Pressure Vessel Components and Heat Leak Material Thickness Diameter Density CD–2012-72 View Port Insulation Wires Structural Standoffs Fused quartz 21.6 cm 4 cm 3 2200 kg/m Aerogel 20 cm N/A 3 20 kg/m Ceramic insulated Ti 21.6 cm 6 mm (including insulation) 3 4500 kg/m Ti alloy (Ti-6Al-4V) 21.6 cm Hollow Tube 5 cm OD, 3 mm thick 3 4430 kg/m 61 March 2012 COMPASS Final Report Thermal Conductivity Quantity Heal Leak In (Total) 1.4 W/mK 2 7.1 W 6.0 COST AND RISK 6.1 Cost 0.017 W/mK NA 108.9 W 21.9 W/mK 24 30.0 W 6.7 W/mK 9 27.8 W Please note that the cost estimates presented in this section should be considered rough order of magnitude (ROM) costs for a S/C that is early in its design phase. In order to estimate the cost of the ALIVE Mission Study S/C design, the MEL generated by the COMPASS team is linked to an Excel-based cost model. Costs are estimated at the subsystem and component levels using mostly mass-based, parametric relationships developed with historical cost data. Quantitative risk analysis is performed on these costs using Monte Carlo simulation based on mass and cost estimating relationship (CER) uncertainties. The pertinent cost modeling assumptions that apply for this S/C design are as follows: § § § § § § § § § § § § The S/C would be designed and built by a prime contractor based on NASA provided specifications. The S/C is assumed to be developed using a proto-flight approach for all subsystems and components. No ground spares are included. Flight heritage is assumed to be OTS for most components as defined by the subsystem leads. However, the electrical power subsystem is assumed to require a new development. The science payload includes the instruments for both the descent science and the surface science as well as the mechanisms for pointing the optical instruments. The cost for these instruments is estimated using the NASA Instrument Cost Model (NICM) for the LIBS, analogies to Galileo for the NMS and ASI, and camera and spectrometer specific CERs for the remaining imagers and spectrometers. The development cost for the Stirling Duplex is based on a CER developed for Non-nuclear Power and Dynamic Isotope Power Systems. The flight hardware for this component is estimated at $20M, based on current estimates for the ASRG which is of similar complexity. The parametric modeling approach assumes that all components are at TRL-6 or higher; therefore, this section does not include any technology development costs necessary to bring any technology up to this level. Software is included as part of a subsystem CER used to estimate the Command and Data Handling subsystem. Planetary systems integration wraps are used to determine costs for Integration, Assembly and Check-out (IACO), Systems Test Operations (STO), Ground Support Equipment hardware (GSE), Systems Integration and Test (SE&I), Program Management (PM) and Launch and Orbital Operations Support (LOOS). The cost estimate represents the ‘most likely’ point estimate based on the cost risk simulation results and roughly equates to the 35th percentile on a pseudo-lognormal distribution. The cost of propellant is not included in these estimates. Costs are in this section are all in FY15$M in order to compare to the New Frontiers cost cap. Taking these assumptions into account, the cost estimate for the COMPASS team S/C design is shown in Table 6.1. The design, development, testing and engineering (DDT&E) represents the non-recurring cost of the S/C while the flight hardware represents the recurring cost. The most-likely cost risk simulation 62 Advanced Lithium Ion Venus Explorer (ALIVE) results for the ALIVE S/C design only (including system integration wraps and prime contractor fee) are shown in Table 6.1 in FY$15M. Table 6.1—COMPASS Subsystem Level Cost Breakdown—ALIVE WBS 06.1.1 06.1.2 06.1.3 06.1.4 06.1.5 06.1.6 06.1.11 Description Lander Science AD&C C&DH Communications and Tracking Electrical Power Subsystem Thermal Control (Non-Propellant) Structures and Mechanisms Aeroshell Cruise Deck Subtotal IACO ST O GSE Hardware SE&I PM LOOS Spacecraft Total (with Integration) Prime Contractor Fee (10% less Science Payload) Spacecraft Total with Fee DDT&E total (FY15$M) 153 46 4 9 10 42 6 35 30 27 209 11 10 20 35 17 14 316 27 343 Flight HW total (FY15$M) 97 34 4 8 10 25 1 15 17 15 129 4 10 20 14 6 14 154 12 166 DD&FH total (FY15$M) 250 81 7 17 21 68 7 50 47 42 339 15 49 23 470 39 508 Figure 6.2 shows lifecycle cost estimate is also included. For this estimate, NASA insight/oversight for the mission is 15% of the prime contractor cost plus fee. Phase A costs are capped at $2.5M per the 2009 New Frontiers AO. The S/C cost represents the development and flight hardware cost from the previous figure. The Mission Operations and Ground Data Systems (GDS) costs consist of a $20M placeholder used to represent the total cost for set-up and operations for this mission. The LV cost is not included in the calculations but assumes an Atlas 411-class LV. Finally, reserves are calculated at 25%. All costs are shown in FY$15M. Table 6.2—Lifecycle Cost Comparison for the ALIVE mission NASA insight/oversight Phase A Spacecraft (with Payload) LV Mission Ops/GDS Reserves Total FY15$M 76 3 508 20 152 760 15% of prime contractor costs NF AO Cost Cap Prime Contractor B/C/D cost plus fee (10% - less science payload) Atlas 411 Mission Ops, GDS, and set-up placeholder cost 25% reserves (less LV) Overall, the mission seems to fit in the higher end of a New Frontiers cost cap, which is assumed to be approximately $775M (FY15$M). However, the reserve posture of this estimate with a minimum 25% reserves would most likely not have enough reserves to be deemed a ‘competitive’ New Frontiers option. But, a stronger reserve posture of 30% or higher exceeds the estimated cost cap. So, this initial analysis shows that the ALIVE mission could potentially compete as a New Frontiers mission in 2015 but would need more reserves to be competitive against other mission proposals. Risk 6.2 Risk Risk Requirements for any S/C design: CD–2012-72 63 March 2012 COMPASS Final Report § The management of risk is a foundational issue in the design, development and extension of technology. Risk management is used to innovate and shape the future § Risk is a chance to do better than planned § Each subsystem was tasked to write a risk statement regarding any concerns, issues and ‘ah ha’s’ § Mitigation plans would focus on recommendations to alleviate, if not eliminate the risk It is important to capture risk in cost estimates, especially technical, schedule and risk data. It may be too early to conduct an in-depth risk analysis, but there are many risks than can and should be identified and addressed at a high level. Cost estimating uncertainty and technical input variable uncertainty need to be considered. In this study, the ground rules and assumptions, data sources, methodology, and the risk assessment are documented to increase credibility and facilitate information sharing, and to make this design/technology usable in the future. Assumptions for any S/C design consist of § § Risk List is not based on trends or criticality Some mitigation plans are offered as suggestions The risk matrix in Figure 6.1 shows a shotgun scatter of where the ALIVE risks are located. Almost all of the risks are considered medium or moderate (yellow) risks. One of the 12 risks were identified as Green or low risks. There was one red risk identified for the X-band science collection system from Venus. All risk owners strive to drive risks and their mitigation steps down to a L×C score of 1×1 within reason. Risks are characterized by the combination of the likelihood (probability) the program element or project will experience an undesired event and the consequences (impact), or severity of the undesired event, were it to occur. In order to establish metrics whereby risks within COMPASS may be assessed on an equitable basis, it is essential that the means for evaluating likelihood and consequences follow the same format. The format is based on a 5×5 risk matrix listed in Figure 6.1. The 5×5 risk matrix contains five adjective ratings for likelihood and five adjective ratings for consequences. Each of the factors (technical, cost, schedule, and safety) must be considered when making a determination of risk Consequence, but a risk need not have impact on all of the four factors. Risk likelihood intends to provide an estimate based on available quantitative data and qualitative experience. Consequence classifications are based on program requirements, project and task performance requirements, mission success criteria, resources, safety, and cost and schedule constraints. Each of the factors (i.e., technical, cost, safety, schedule) must be considered when making a determination of risk consequence. Figure 6.2 to Figure 6.7 describe the risk statement, the risk context, and possible mitigation plans for each risk identified. 64 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 6.1—ALIVE Risk List. Figure 6.2—Risks 1 and 2—Mission and Power risks for ALIVE. CD–2012-72 65 March 2012 COMPASS Final Report Figure 6.3—Risks 3 and 4—Mechanisms and Mission risks for ALIVE. Figure 6.4—Risks 5 and 6—Thermal and Mission risks for ALIVE. 66 Advanced Lithium Ion Venus Explorer (ALIVE) Figure 6.5—Risks 7 and 8—Electronics and Structures risks for ALIVE. Figure 6.6—Risks 9 and 10—Thermal risks for ALIVE. CD–2012-72 67 March 2012 COMPASS Final Report Figure 6.7—Risks 11 and 12—Thermal and Propulsion risks for ALIVE. Figure 6.8—ALIVE TRL assessment. 68 Advanced Lithium Ion Venus Explorer (ALIVE) 7.0 BIBLIOGRAPHY System AIAA S-120-2006, AIAA Standard Mass Properties Control for Space Systems. ANSI/AIAA R-020A-1999, Recommended Practice for Mass Properties Control for Satellites, Missiles, and Launch Vehicles. Larson, W.J. and Wertz, J.R. (eds.), (1999) Space Mission Analysis and Design, Third Edition, Space Technology Library, Microcosm Press. Mission Anderson, David J., Pencil, Eric J., Liou, Larry C., Dankanich, John W., Munk, Michelle M., and Hahne, David, (2010), “The NASA In-Space Propulsion Technology Project’s Current Products and Future Directions,” AIAA–2010–6648, 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, July 25–28, 2010, Nashville, TN. Balint, Tibor, (2002),”Summary of Russian Planetary Lander Missions.” Deep Space Mission Architecture Group, NASA-JPL, CA “COMPASS Final Report: Mars Earth Return Vehicle (MERV),” CD–2009–31, February–March 2009. Desai, P.N., Braun, R.D., Engelund, W.C., Cheatwood, F.M., Kangas J.A., (1998), “Mars Ascent Vehicle Flight Analysis,” 7th AIAA/ASME Joint Thermophysics and Heat Transfer Conference June 15–18, 1998, Albuquerque, NM, AIAA 98-2850. Dillman, Robert and Corliss, James, (2008), “Overview of the Mars Sample Return Earth Entry Vehicle,” Sixth International Planetary Probe Workshop, June 26, 2008. Dux Ian J., Huwaldt, Joseph A., McKamey, R. Steve, Dankanich, John W., (2010) Mars Ascent Vehicle Gross Liftoff Mass Sensitivities for Robotic Mars Sample Return, NASA Deep Space Missions, NASA TM 2010. ESA Aurora Program, http://www.esa.int/esaMI/Aurora/SEMCWB1A6BD_0.html Holmberg, Neil A.; Robert P. Faust, H. Milton Holt (1980), “Viking ‘75 Spacecraft Design and Test Summary Volume III– Engineering Test Summary,” NASA Reference Publication 1027, Nov. 1980. Johnston, M.D., Esposito, P.B., Alwar, V., Demcak, S.W., Graat, E.J., and Mase, R.A. (1998), “Mars Global Surveyor Aerobraking at Mars,” AAS 98-112 http://Mars.jpl.nasa.gov/mgs/sci/aerobrake/SFMech.html Matousek, S., Adler, M., Lee, W., Miller, S.L., Weinstein, S., (1998), “A Few Good Rocks: The Mars Sample Return Mission Architecture,” AIAA/AAS Astrodynamics Specialist Conference, August 10–12, 1998, Boston, MA, AIAA–98–4282. Palaszewski, B. and Frisbee, R., (1988), “Advanced Propulsion for the Mars Rover Sample Return Mission,” AIAA–88–2900, AIAA/ ASME/SAE/ASEE 24th Joint Propulsion Conference, July 11–13, 1988, Boston, MA Preliminary Planning for an International Mars Sample Return Mission, Report of the International Mars Architecture for the Return of Samples (iMARS) Working Group, June 1, 2008. Rose, J., (1989), “Conceptual Design of the Mars Rover Sample Return System,” 27th Aerospace Sciences Meeting, Jan. 9–12, 1989, Reno, NV, AIAA–89–0418. Spencer, D.A., Tolson, R., (2007), “Aerobraking Cost and Design Considerations,” Journal of Spacecraft and Rockets, Vol. 44, No. 6, Nov/Dec. 2007 Stephenson, David, (2002), “Mars Ascent Vehicle—Concept Development,” 38th Joint Propulsion Conference and Exhibit, July 7–10, 2002, Indianapolis, IN, AIAA 2002-4318. Whitehead, John C., (1997), “Mars Ascent Propulsion Options for Small Sample Return Vehicles.” 33rd AIAA/ASME/SAE/ ASEE Joint Propulsion Conference and Exhibit, July 6–9, 1997, Seattle, WA, AIAA 97-2950. Williams, R., Gao, Y., Kluever, C.A., Cupples, M., and Belcher, J., (2005), “Interplanetary Sample Return Missions Using Radioisotope Electric Propulsion,” 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit , July 10–13, 2005, Tucson, AZ, AIAA–2005–4273. Zubrin, Robert, (1996), “A comparison of approaches for the Mars Sample Return Mission,” AIAA 34th Aerospace Sciences Meeting and Exhibit, Reno, NV, Jan. 15–18, 1996, A9618450, AIAA Paper 96-0489. GN&C L. Norris, Y.C. Tao, R. Hall, J. Chuang, M. Whorton, “Analysis of Ares I Ascent Navigation Options”, AIAA 2008-6290. Structures D. Persons, L. Mosher, T. Hartka, (2000), “The NEAR and MESSENGER Spacecraft: Two Approaches to Structure and Propulsion Design,” AIAA–00–I406, A00-24531. Zibdeh, Hazim S., and Heller, Robert A., (1989), “Rocket motor Service Life Calculations Based on the First-Passage Method,” Virginia Polytechnic Institute and State University, Blacksburg, VA, Journal of Spacecraft and Rockets, 0022-4650, Vol. 26, No. 4 pp. 279–284. Thermal Aerodynamic Heating Wikipedia entry. http://en.wikipedia.org/wiki/Aerodynamic_heating Alexander, M. (ed.) (2001), “Mars Transportation Environment Definition Document,” NASA/TM—2001-210935. Bejan, A. and Kraus, A.D., (2003), Heat Transfer Handbook, John Wiley & Sons. CD–2012-72 69 March 2012 COMPASS Final Report Chapman, A.J., (1974), Heat Transfer, Third Edition, Macmillan Publishing Company. Gilmore, David G. (ed.), (2002), Spacecraft Thermal Control Handbook: Volume 1 Fundamental Technologies, AIAA. Hyder, A.J., Wiley, R.L., Halpert, G., Flood, D.J. and Sabripour, S., (2000), Spacecraft Power Technologies, Imperial College Press. Incopera, F.P. and DeWitt, D.P., (1990), Fundamentals of Heat and Mass Transfer, John Wiley and Sons. Larson, W.J. and Wertz, J.R. (eds.), (1999), Space Mission Analysis and Design, Third Edition, Space Technology Library, Microcosm Press. Olds, J. and Walberg, G., (1993), “Multidisciplinary Design of Rocket-Based Combined Cycle SSTO Launch Vehicle using Taguchi Methods,” AIAA 93-1096, AIAA/AHS/ASEE Aerospace Design Conference, February 16-19, 1993, Irvine, CA. Penuela, David, Simon, Mathew, Bemis, Eammon, Hough, Steven, Zaleski, Kristina, Jefferies, Sharon and Winski, Rick, “Investigation of Possible Heliocentric Orbiter Applications for Crewed Mars Missions,” NIA/Georgia Institute of Technology, Project 1004. http://www.nianet.org/rascal/forum2006/presentations/1004_nia_paper.pdf RP–07–100_05–019–I; Volume I: Ice Mitigation Approaches for Space Shuttle External Tank Final Report. Sutton, Kenneth and Graves, Randolph A. Jr., (1971), “A General Stagnation Point Convective-Heating Equation for Arbitrary Gas Mixtures,” NASA TR R-376. http://kids.britannica.com/comptons/art-94135/A-surface-map-of-Venus-shows-the-planets-global-topography http://www.sageofathens.com/Documents/Duplex.pdf 70 Advanced Lithium Ion Venus Explorer (ALIVE) APPENDIX A—ACRONYMS AND ABBREVIATONS ACS AD&C AIAA Al ALIVE ANSI AO ASC ASI ASRG AWG C&DH C&T C&TN CAM CBE CEA CER CO2 Comm COMPASS COTS DCIU DD&FH DDT&E DPU DSN DTE DTU Eb/N0 EDL CD–2012-72 EDS EEE EIRP ELV EMP EP EPC EPOXI EVE FOM FPGA FY GDS GLIDE Attitude Control System Attitude, Determination & Control American Institute for Aeronautics and Astronautics aluminum Advanced Long-Life Lander Investigating the Venus Environment American National Standards Institute Announcement of Opportunity Advanced Stellar Compass Atmospheric Structure Investigation Advanced Stirling Radioisotope Generators American Wire Gauge Command and Data Handling command and telemetry Communications & Tracking Network collision avoidance maneuver current best estimate GN&C GRC GSE HAN HGA IACO IR Isp JPL KSC LGA LGA Li LIBS cost estimating relationships carbon dioxide communications COlaborative Modeling and Parametric Assessment of Space Systems commercial off the shelf Digital Control and Interface Unit? LiDS LNT LOCO design, development, test, and evaluation data processing unit Deep Space Network data terminal equipment Technical University of Denmark energy per bit to noise power spectral density ratio entry, descent, and landing LOOS LPF LRO LV 71 Earth departure stage equivalent isotropic radiated power expendable launch vehicle electromagnetic pump electric propulsion ? Extrasolar Planet Observation Extended Venus Explorer figure(s) of merit Field Programmable Gate Array fiscal year Ground Data Systems GLobal Integrated Design Environment Guidance, Navigation and Control NASA Glenn Research Center Ground Support Equipment hydroxyl-ammonium nitrate high gain antenna Integration, Assembly and CheckOut infrared specific impulse NASA Jet Propulsion Laboratory NASA Kennedy Space Center low gain antenna lunar gravity assist lithium Raman/Laser Induced Breakdown Spectroscopy Lithium Duplex Sterling lithium nitrate trihydrate LOw COmplexity LOssless COmpression Launch and Orbital Operations Support pg18 Lunar Reconnaissance Orbiter launch vehicle March 2012 COMPASS Final Report MAC MALTO MASTIF MEL MER MET Mg MGA MLI MMPDS MOP MSL N/A NaK NaS NASA Nav NIAC NICM NIST NMS OTS OU P&ID PAF Pan Cam PEL PICA PLA PM PMAD PN PV R&D3 media access control Mission Analysis Low-Thrust Optimization Mission Analysis and Simulation Tool In Fortran Master Equipment List Mars Exploration Rover pg12 magnesium mass growth allowance multilayer insulation Metallic Materials Properties Development and Standardization RAM RCS RF RFI ROM S/C SA SDI SDO SDST SE&I SEU SIRU Mars Science Laboratory not applicable sodium-potassium alloy sodium-sulfur National Aeronautics and Space Administration navigation NASA Innovative Advanced Concepts NASA Instrument Cost Model National Institute of Standards and Technology Neutral Mass Spectrometer off-the-shelf optical units SLOC SOAP STO SUA TBD TBR TCM TCS TDRSS TFDoM Ti TLI TLS TRL TT&C TWTA VFDRM VITaL WBS WGA WGS payload attach fitting Panoramic Camera Power Equipment List Phenolic Impregnated Carbon Ablator Payload Adaptor Program Management power management and distribution pseudo-noise photovoltaics 72 Research and Development Degree of Difficulty random access memory Reaction Control System radio frequency radio frequency interference rough order of magnitude spacecraft solar array serial digital interface serial data output Systems Integration and Test single event upset Scalable Inertial Measurement Unit source lines of code Satellite Orbit Analysis Program Systems Test Operations systems uncertainty analysis to be determined to be resolved trajectory correction maneuvers Thermal Control System Tracking and Data Relay Satellite System Test Facility Degree of Modification titanium trans-lunar injection Tunable Laser Spectrometer technology readiness level telemetry, tracking and command traveling wave tube amplifier Venus Intrepid Tessera Lander work breakdown structure weight growth allowance weight growth schedule Advanced Lithium Ion Venus Explorer (ALIVE) APPENDIX B—RENDERED IMAGES B.1 Insert subtitle Figure B.1— Figure B.2— Figure B.3— CD–2012-72 73 March 2012 COMPASS Final Report APPENDIX C—COMPASS INTERNAL DETAILS (ALWAYS LAST) C.1 COMPASS Description The COncurrent Multidisciplinary Preliminary Assessment of Space Systems (COMPASS) team is a collaborative engineering team whose primary purpose is to perform integrated-vehicle systems analysis and provide trades and designs for both Exploration and Space Science Missions. C.2 GLIDE Study Share GLobal Integrated Design Environment (GLIDE) is a data collaboration tool that enables secure transfer of data between a virtually unlimited number of sites from anywhere in the world. GLIDE is the primary tool used by the COMPASS design team to pass data real between subsystem leads in real-time. While GLIDE 2 was being tested during this design session, the old shares are being used to store the data and the MELs. The data on the share can be found here: https://glidesharename/XXXX C.2.1 GLIDE Architecture For this study, the COMPASS Team is testing the GLIDE 2 application and server. The architecture and database information will be referencing the GLIDE 2 server. Architecture: C.2.2 XXX GLIDE Study Container Table C.1 lists the study container and descriptions of the cases run with the GLIDE-specific data necessary for the COMPASS Team members to conduct the study. Table C.1—Study Container and Descriptions Study name Case no. Description 74 Study container XXX_Case0