H ybrid P ropellant M odule H ybrid P ropellant M odule Block 2 Update 12/2/2001 Pat Troutman LaRC Spacecraft & Sensors Branch [email protected]

HybridPropellantModule

HybridPropellantModule

Block 2 Update 12/2/2001

Pat TroutmanLaRC Spacecraft & Sensors [email protected]

Block 2 Update 12/2/2001

Pat TroutmanLaRC Spacecraft & Sensors [email protected]

Future Assumptions: 2015 and Beyond

Low Earth Orbit (LEO) & Beyond: • NASA/International Space Exploration

• NASA has deployed a gateway facility at the Earth-Moon L1 point.

• ISS has evolved into a transportation hub & servicing facility.

• Commercial• Commercially viable in-space manufacturing of pharmaceuticals and materials resulting from ISS research has begun on automated and crew tended platforms

• A commercially owned upgraded Shuttle features a payload bay passenger module for commercial crews and other paying passengers.

• The first hotel in space (based on the NASA gateway facility and catering to the elite) has opened in LEO.

• Military•The United States military dominates the space theatre.

Available Earth-to-Orbit Transportation: • Upgraded Shuttle - operations overhead cut in half with the same performance.

• Large reliable ELV - 35,000 kg to LEO with a 6 meter shroud.

• Inexpensive ELV - weekly launch of 10,000 kg of logistics to LEO.

• Revolutionary RLV eventually replaces weekly ELV launches.

HPM Level 1 Requirements

• The HPM shall support commercial, NASA and DOD missions.

• The HPM and associated elements shall be designed, built, inspected, tested, and certified specifically addressing the requirements for human-rating.

• The HPM shall be reusable

• The HPM shall be designed for an operational lifespan of ten years.

• The HPM shall have on-orbit maintainable (via EVA and robotics) avionics.

• The HPM shall be capable of autonomous operations for all systems

• The HPM shall accommodate automated rendezvous and docking with other vehicles.

• The HPM shall provide long-term storage of Lox, LH2, & Xenon for use by chemical and electric propulsion systems

• The HPM shall be capable of being refueled on-orbit

HPM Deployed Position (Dia = 4.8m)

Deployed Docking/Fluid Transfer Ring (2)

HPM Configuration

HPM Upper Stowed (Dia = 4.5m)

PV Arrays in Stowed Position (44m2 total)

Trunnion Fittings (4)

Grapple Fixture

ORUs

Stowed

Deployed

Access Hatches

Stowed Docking/Fluid Transfer Ring (2)

Tank Supports

Radiators (2)

HPM Interior Layout

HPM Structural Layout with Tanks

LH2 Tank Properties:

Volume = 65.8m3

Chemical Mass = 4455 kg

Tank Mass = 162 kg

LOX Tank Properties:

Volume = 24.19m3

Chemical Mass = 26,723 kg

Tank Mass = 48 kg

Xe Tank Properties:

Volume = 3.85m3

Electric Mass = 13,552 kg

Tank Mass = 10 kg

HPM Overall Dimensions & Capacities

6.35m7.75m

1m 3.15m

Once deployed the docking rings can extend to 0.5m

A

A

B

B

Radiators

Lower debris shielding

LOX tank walls and insulation Place holders

for radiation and thermal protection

Keel fitting

Lower HPM Cross Section

The lower section of the HPM uses the Whipple debris shielding similar to existing portions of the ISS. This type of shielding incorporates a constant outer diameter allowing access to the subsystems.

The lower primary structure consists of the inner most wall of the debris shielding tied into the stringers that run the length of the lower section. The vacant space will be filled with thermal protection layers.

Upper section deployed debris shield

I-Beams

LOX Tank Wall and Insulation (2.54cm thk)

Radiation Shielding Place Holder

Upper HPM Cross Section

The upper section of the HPM uses the deployable debris shielding similar to that demonstrated on the Mars Trans Hab Module. Although heavier than the simple Whipple type shielding, it can withstand impacts from much larger diameter micrometeoroids and orbital debris. The upper structure is similar to the lower section but just at a larger diameter.

A A

Section AA

LH2 Tank Properties:

Volume = 65.8m3

Surface Area = 86.0m2

Barrel Length = 4.44m

Inner Diameter = 3.68m

LOX Tank Properties:

Volume = 24.19m3


Barrel Length = 1.27m

Inner Diameter = 3.30m

Intl. Berthing Docking Mechanism (IBDM)1 (2)Max Dim’s: 1.4m dia x 0.25m thickHatch Pass Through: 0.80m

1IBDM in development, estimated year 2005 operational date

PV Drive Location (2)

Avionics ORUs

Flywheels

Cryogenic Coolers (2) – The other Cooler is located between the LH2 and LOX Tank

Fluid Transfer Line Routing

Xe Tank Properties:

Volume = 3.85m3


Upper Deployed Debris Shield (Dia = 4.8m - 0.305m thick)Y

PV Array Area = 12m2 per side

Radiators (2)

Lower Debris Shield (0.101m thick)

Tank Supports (Similar for LOX tank)

Supporting Structure (0.305m I-Beams)

FTI

14m

HPM Structures Technology Summary

Current Technology Research Activities

TechnologySummary Description of Desired

Technology and Key Performance MetricsCurrent

TRLWhere Who

Current Funding

(K$)

Increase in Funding

Required (no, small, major)

Applications of the Technology Other

than HPM

Carbon Composites High stiffness to low weight ratio 4 various various 12,824 none various

Multi-function StructureSecondary structure of all subsystems incorporated into primary structure

5 various various 5,394 none various

Ceramic Matrix Composites

High stiffness to low weight ratio and high thermal protection


Metal Matrix Composites

High stiffness to low weight ratio and MOD protection 3 various various 3,877 small various

Graphitic FoamLightweight filler for debris shielding that also adds thermal protection

5 various various 300 small various

Syntactic Metal Foam High stiffness to low weight ratio 4 various various 70 small various

Micro-Meteoroid & Orbital Debris Shielding

Technology to mitigate the threat of meteoroid damage 3 various various 500 major various

Insulation MaterialsThermal protection from space environment and cryogenic thermal stability

2 various various 140 major various

Self Healing MaterialsAbility to repair pressure/structure walls after debris impact - TBD on HPM

2 various various 70 small various

Biomimetic MaterialsMimicking structures found in nature to help reduce loads and stress concentrations - TBD on HPM


Carbon Nanotubes High stiffness to low weight ratio - TBD on HPM 2 various various 5,539 none various

HPM Systems

HPM Guidance, Navigation & Control System

Functional Description:Maintain attitude in free flight mode and during autonomous docking operations in LEO,GEO and L1 environments.

Key Performance Requirements:Hold attitude to within +/- 5 degrees of TEA during LEO/GEO parking orbit modePosition and hold attitude to within +/- 0.5 degrees during docking operationsProvide attitude and position knowledge in support of automated docking operations

Design/Technology Description:

Attitude Control – Flywheels used to rotate and maintain attitude. The flywheels are also integrated with the power system as an energy storage device.

Position & Attitude Knowledge - The attitude, attitude rates, position, and velocity, and Sun pointing of the HPM would be determined using an enhanced Microcosm Autonomous Navigation System (MANS) sensor suite, comprising of Star Sensor and Earth sensor with IMU as back-up. MANS suite can currently provide 100m position information, 0.03 deg attitude information, and is light and uses little power. While MANS has been used for Earth orbits, it’s extension to deep space applications is new technology.

Autonomous Rendezvous & Docking - The Autonomous Formation Flying (AFF) sensor would be used by the HPM and other docking vehicles for precision relative navigation during automatic rendezvous and docking. The AFF can provide 1 cm position accuracy, 0.1mm/s relative velocity, 1 arc-minute attitude, using 1W power and weighs less than 2 kg. This technology needs to be demonstrated on-orbit. AFF can replace or enhance GPS and retro-reflector based concepts.

HPM Guidance, Navigation & Control (Schematic)

HPM Attitude Dynamics

HPM Translational Dynamics

MANS

Star Camera Earth Sensor

IMU

HPM Attitude Controller and momentum manager

Communications subsystem

To Earth/Moon or other vehicles for coarse nav information

Attitude and attitude rates

Position and velocity

Flywheel

Flywheel momentum

Flywheel torque

Steering Law

Flywheel torque command

Power subsystem

AFF sensors

Docking Vehicle dynamics

Docking vehicle orbit control

Power profile

Wheel speeds

Thrust

Relative nav

HPM - Guidance, Navigation & Control (Technology Requirements)

Technology Description/

Metrics

TRL Current Technology Research Activities Other Applications

Where Who Funding Increase?

MANS Microcosm

Autonomous Navigation

System

Requires development of software, hardware definitions/interfaces, testing for deep space platforms.

5 Micro-cosm Inc.

Gwynne Gurevich

Phone: (310) 726-4100

TBD

Can use NASA SBIR

Small Attitude and Position info for any satellite in space, near Earth or deep space

AFF Autonomous

Formation Flying

Based on GPS technology and can work in deep space with or without GPS satellites. Needs on-board implementation & testing.

• 1cm relative position

• 0.1mm/s relative velocity

• 1arc-minute attitude

• Average Power=1W

• Less than 2kg

3 JPL Kenneth Lau

Initial funding

from NASA

complete

Small For any

in-space rendezvous and docking between spacecraft & formation flying.

•100m position info• 0.03deg attitude info • 11kg & 28W based on sensor suite used

C&DH/C&T Systems & Technologies

Functional Description:Communications, command, telemetry, and recorder systems

Key Performance Requirements:

•Low data rate: 2000 bps or less telemetry,1000 bps or less command

•1 day data storage capability•3 dB link margin from E-S L2

•Low gain patch antennas•Medium gain horn antennas•Redundancy in all three systems

M - Moderate technology mass = 22 KgE - Extreme technology mass = 8 Kg

Technology (C&T) Mass PowerCurrent capability 31 Kg 65 WShrink power amp/transponder 20 45Shrink power amp/transponder 10 35

Technology (C&DH) Mass PowerCurrent capability 11 Kg 39 WIntegrated system 7 25System on a Chip 3 15

Key Technologies

C&DH/C&T

Attached vehicles(CTM, SEP)

S-Band Communications System

Computers RecordersOther systems

Power, Prop, etc.

Attached vehicles(CTV, Gateway, OTV, ISS)

Power Amp

Transfer

SW

Power Amp C&DH

Diplexer

DiplexerRF SW

RF SW

S-Band Transponder

S-Band Transponder

C&DH

SW

Transfer

C&DH

S-Band Communications (C&T)




TRLWhere Who

Current Funding

(K$)

Increase in Funding


Other Applications of the Technology Other than HPM

Integrated systemIntegration of computer/data storage systems. Mass & power none Small Satellites

System on a chip Move majority of functions into single chip. Mass & power Small Satellites

Shrink power amp & transponder

Reduce mass and power for power amps and transponders Small Satellites

HPM C&DH/C&T Technology Summary

Propellant Management System & Technologies

Functional Description:Efficient systems for transfer and storage of cryogenic fluids for long periods of time

Key Performance Requirements:• LH2 Tank Volume = 65 m**3• LO2 Tank Volume = 24.2 m**3• LXe Tank Volume = 3 m**3• Loss kg/month = near zero• Re-usability: 4 fill and drain cycles per

year for 10 years with no refurbishment

Design/Technology Description:

Take advantage of the tremendous advances in cryo-cooler technology and combine active (cryo coolers) and passive (multi-layer insulation-MLI) thermal control technologies to remove heat entering a cryogenic propellant tank and control tank pressure.

Develop new technology to routinely and autonomously transfer cryogenic propellants for in-space operations.

Cryocooler

ColdFinger

MLI Blankets

Vapor

Cryogen

Space

Heat Exchanger

RadiatorSolarArray

Possible ZBO In-Space Configuration

Propellant Management System

LOX

LH2

H2 Vent

Pressure Building

Coil

Pressure Building

Coil

O2 Vent

Xe Xe

Xe VentXe Vent

Key

Valve

Relief Valve

Burst Disk

Can we put the Cryo-coolers on this schematic?



TRLWhere Who

Current Funding

(K$)

Increase in Funding



than HPM

Lightweight Tank Materials

Use Composite Materials to halve the weight of propellant tanks 6

MSFC 3000 small

Space Launch Initive

Lightweight Componet Materials

Use Composite Materials to halve the weight of lines, valves and fittings 3

GRC 250 major

Space Launch Initive

Lightweight Docking Adaptors

Use Modern design and techniques to reduce the weight of docking adaptors 3

JSC ?? major

Crew Transfer Vehicles

Cryogenic Transfer Efficienly transfer large quantities of cryogenic liquids in low gravity 4

GRC 0 major

All deep space missions

Long Life Cryocoolers for Zero Boiloff

Develop highly reliable long-life cryocoolers to remove thermal energy for long term storage 4

ARC 500 small

Sensor Cooling

Long Life Valving Develop long-life eletric acuated valves with low sealing forces and seat wear capable of functioning at cryogenic temperatures with miniumal leakage

3

GRC 0 major

All deep space missions

HPM Propellant Management Technology Summary

POWER GENERATION Specifc Power1 EfficiencyMBG2 Crystalline PV 200 W/kg 30%Thin Film PV 200 W/kg 10%MBG Crystalline PV 250 W/kg 40%Thin Film PV 270 W/kg 15%Thin Film PV 600 W/kg 20%Advanced Array Designs >400 W/kg >40%Quantum Dots3 >500 W/kg 60%

Functional Description:The Electrical Power System consists of power generation for HPM (house-keeping), CTM, CTV and the SEP Stage, energy storage for HPM, CTM, and CTV power during shadow, and power processing.

Key Performance Requirements:• Minimal system mass and volume• Reliability; cycling capability• Radiation degradation resistant for system lifetime of 10 years• Capable of power generation with arrays stowed (at reduced level)• Redeployable

ENERGY STORAGE Specific Cycle Depth ofEnergy Lifetime/ Discharge

EfficiencyLi-based batteries4 100 Wh/kg 30 kCyc. 60%Century Flywheel5 45 Wh/kg 75 kCyc. 89%Active Dedicated RFC6 400 Wh/kg 55% Eff.Li-based batteries4 200 Wh/kg 30 kCyc. 70%Advanced Flywheel5 100 Wh/kg 75 kCyc. 89%Passive Unitized RFC6 1000 Wh/kg 65% Eff.Full polymer batteries4 300 Wh/kg 20 yrs (GEO) 70%Future Flywheel5 150 Wh/kg >95 kCyc. 90%Passive Unitized RFC6 >1000 Wh/kg 80% Eff.

POWER PROCESSING Specific Efficiency TemperatureEnergy

Converter w/Active Control 0.5 kW/kg 90% 125 °C300V Power Distribution 0.3 kW/kgModular, High-Temp. 1.5 kW/kg 95% 225 °C Converters600V Power Distribution 0.7 kW/kgHigh-Temp. PMAD System 3.0 kW/kg 95% 350°C1200V Power Distribution 1.4 kW/kg

Notes:1 - Array level specific power2 - Multiple band gap cells (I.e. 2, 3, and 4 junctions)3 - High Risk/high Potential technology4 - Does not include power electronics mass5 - Includes power electronics mass6 - Regenerative Fuel Cell: Specific Energy is a function of discharge time

Power System

Hybrid Propellant Module Power System

Performance GuidelinesPower Generation:3.1 kW required at 100% duty

cycleDuring LEO to Earth-Moon

L1 transfer and returnEnergy Storage:Provide required power during

shadow/chemical rocket useCross-use capability

Operating Life:10 year system lifetime (~ 5

round trips)

Preliminary ConfigurationPV Arrays:

Rigid Planar structuresStowed on exterior of HPMArrays retracted during

chemical rocket firingEnergy Storage:

Flywheel systems5.1 kWh capacity requiredSharing functions with

attitude control system

Flywheels

Solar Array

Charge/Discharge

System

Thermal control

Solar Array

Power ProcessingPower Regulation & Control

Power Distribution

To Spacecraft Bus

Baseline EPS Schematic


RESULTS

Technologies selected:Advanced crystalline multi-band gap

photovoltaics47% eff. @ AM0, 135 W/kg @ panel

Flywheel energy storage5.1 kWh capacity, 3.3 kW delivered, 89% DOD

Advanced power processing95% eff., 670 W/kg

PV Arrays 64Component

MassArray Structure: drives, yoke,

mechanisms, harness46

Breakdown Energy Storage 136Power Processing 15

Thermal Control System 3Total EPS Mass, kg 264

EPS Mass, kg at 2016 level 293Array Area, m^2 20.2

System Wing Dimensions, m 3.2 x 3.2Characteristics Power Generated @ BOL, kWe 10.1

Energy Storage Capacity, kWh 5.1

Mass Breakdown

Power Processing6%

Thermal Control System

1%PV Arrays

24%

Array Structure: drives, yoke, mechanisms,

harness17%

Energy Storage52%


Technology Requirementsmy preliminary guesses - will fill out with better info



TRLWhere Who

Current Funding

(K$)

Increase in Funding



than HPM

PhotovoltaicsHigh Efficiency Multi-band gap cell material, 40% efficiency expected 15 years out (>40% targeted)

2/5GRC,

Industry? Major

All spacecraft applications

FlywheelsComposite wheels with lightweight power electronics and

containment housing, capable of providing momentum control for s/c

2GRC,

Industry? Major

Long-duration spacecraft

applications

BatteriesLithium-based batteries, >200 Wh/kg, >30 kCyc., 70%

DOD at GEO4 Industry ?

All spacecraft applications

Power ProcessingLightweight power conversion and switching electronics,

> 1 kW/ kg for distribution, > 2 kW/kg for conversion, capable of high temperature operation

4?GRC, Lots of

Industry? Major

All space systems; many terrestrial

applications

Thermal Control Lightweight radiator materials, 4 kg/m 2̂, operation at

high temperatures ? ?

All temperature-sensivite space

systems


Structures20.8%

Shielding40.6%

Command/Control/Comm1.1%Navigation/

Attitude Control0.3%

Pow er6.8%

Thermal2.4%

Propellant Managment

28.0%

Mass Properties – HPM Block 2

Subsystem Calculated Mass (kg)

Navigation/Attitude Control 11.84

Command/Control/Comm 41.50

Thermal 93.35

Power 264.00

Propellant Management 1,089.44

Structures 1,240.00

Shielding 1,582.00

Calculated Dry Mass 4,322.13

Dry Mass Margin -218.13

Dry Mass Target Mass 4,104.00

Documents

H ybrid P ropellant M odule H ybrid P ropellant M odule Block 2 Update 12/2/2001 Pat Troutman LaRC Spacecraft & Sensors Branch [email protected]