National Aeronautics and Space Administration
Small Stirling Technology Exploration Power for Future Space Science Missions
Scott Wilson, Nicholas Schifer, NASA Glenn Research Center
March 4, 20192019 IEEE Aerospace Conference
2019 IEEE Aerospace Conference
National Aeronautics and Space AdministrationNational Aeronautics and Space Administration
Michael Casciani, Vantage Partners, LLC.
National Aeronautics and Space Administration
Small nuclear power systems could provide electricity to power probes, landers, rovers, or communication repeaters for space science and exploration missions• Would convert heat to electricity for powering spacecraft sensors and communication
• Fractional GPHS (General Purpose Heat Source) output: ~60 watts thermal output (good for 10-Watt Generator)
• LWRHU (Light Weight Radioisotope Heater Unit, often called RHU) ~1 watt thermal output for each heater unit
Low Power RPS for Small Spacecraft
GRC Development Goals• Sufficient power for small spacecraft functions
• Long-life and low degradation to ensure sufficient power at EOM
• Robust to critical environments (vibration, shock, constantacceleration, radiation, etc.)
• Thermal capability and high efficiency
• Operates in vacuum or atmosphere
Dynamic Power Conversion System Efficiency• 1 watt generator efficiency projected to be 12-13%
• 10 watt generator efficiency projected to be 16% Conceptualization of Seismic MonitoringStations Being Deployed from Rover
JPL Pub 04-10, Sept-2004: 51 mission concepts, 14 RHU-based, 13 fractional-GPH
2
National Aeronautics and Space Administration
• Subassemblies under development include convertor, controller, and insulation
• Need to formalize requirements later this year (Fundamental Research Project to help)
• Goals are based on existing DPC requirements
GRC 1-W Dynamic RPS Concept - Goals
Category Goals Current Best Estimate
Design life 20 years 20 years
Heat input 7 to 8 watts to convertor 8 watts
Power output At least 1 We DC from controller > 1 We DC
Heat source surface temperature TBD < 450 ºC
Stirling hot-side temperature 325 to 375 ºC 350 ºC
Stirling cold-side temperature -100 to 50 ºC -100 to 50 ºC
Robustness Overstroke tolerant
Random vibration level TBD
Environment vacuum or atmosphere vacuum and atmosphere
Constant acceleration 20 g 19 g
3
National Aeronautics and Space Administration
In-house dynamic RPS from 1-10 We DC power output• Demonstrate practicality of 1 watt power level by maturing subassemblies
and interfaces
• Complete engine design at 10 watts power output (62 Wth heat source)
Initial Demonstration • Controller breadboard• Free-piston Stirling convertor• Multi-layered metal foil insulation (using Stirling thermal simulator)
Increased Fidelity • Controller brassboard fabrication and test• Stirling convertor durability testing
System Testing • Stirling convertor + controller• Electrically heated prototype system (includes insulation)
Technology Development at GRC
4
Pneumatic Testing
Controller Testing
Insulation Testing (at vendor)
National Aeronautics and Space Administration
GRC 1-W Dynamic RPS Concept
Electrical Controller
Stirling Engine
Multi-Layer Metal Insulation
Heat Source
Linear Alternator
Stirling Convertor
Radiative Coupling Heat Rejection Flange
5
in-line concept for Stirling convertor not yet complete
Table 1: smal/STEP Concept Characteristics ,--------------------------------------------------1 ' ' ' ' ' ' ' ' ' ' ' ' ' ' ' ' [ _ - - - - - - - - - - - - ----- - - - - - - - - - - - - --- - - - - - - - - - - - - ----- I
Item smallSTEP
Nuclear Fuel 8 LWRHU
Thermal Inventory 8 watts (thermal) Beginning of Life ,(SOL) -when the RPS is fueled
Beginning of Mission 1 watt (BOM) Power - up to 3
vears after fuelinq End Of Design Life 0.83 watts
(EODL) Power - 14 years after BOM
Power Conversion Stirling Converter
Environment Multi-Miss ion Capable
Voltage Range 4 to 5 V DC
Degradation Rate 1.2%/yr
Efficiency 13%
Mass (including 6 pounds (3 controller) kilograms)
Volume 4.3 inch (11 centimeter) X 12.5 inch (32
centimeter) cube
National Aeronautics and Space Administration
Light Weight Radioisotope Heater Unit (LWRHU)• Long history of use on many space missions for
heating spacecraft electronics
• Aeroshell designed to survive reentry into Earth’s atmosphere for safety
• Diameter: 1.0 inch, Length: 1.3 inch (1.1 watts of thermal energy at BOL)
Generator concept uses 8x LWRHU• 8 Wth Heat to 1 We DC electrical power
GRC testing will use electric heaters to simulate the LWRHUs• Designed to provide similar thermal gradients
compared to LWRHU
• There are four resistance cartridge heater total, each one simulates two LWRHUs
Heat Source
LWRHU Assembly
LWRHU Simulator uses electric heaters
Welded wire bus
LWRHU volume
electricheater
cross-section view Test Hardware
Graphite block
6
Generator concept uses 8x LWRHU
o • •
National Aeronautics and Space Administration
Objectives • High performance is critical to minimize losses
• Peregrine Falcon Corp. provided Multi-Layered Metal Insulation (MLMI) package in October 2018
• MLI design considerations
• 8 watts thermal input, targeted >90% efficient
• Evacuated to enable low thermal losses
• Lab unit allows disassembly for inspection
• Thermal simulator design considerations
• Reject 7 watts
• Radiation coupling to heat source
• Instrumented with TCs along the length to calculate heat transfer, also enables model comparison
• Efficiency is flat with achievable emissivity values
Insulation Design
7
MLI Package
Thermal Analysis Study
Thermal Simulator Thermal SimulatorFabricated
J ~O
,,.
2,0
liO
.,
1.)
f' 1.0
-~ 0.8
~ 0.6 C
~ OA ;;; C 0 ,2
00
111
00
0
•
-300
Varying Fm issivity
rel;;ati'-'fl'Yfl~
~ ·~- .. Harder
0.0~ 0 .1
• • Easler
0 .15
Environm~nt al Te mperature, C
Varyi ne [nvironme nta l Tem perat ure
deep space
• e~rth
02
• Lunar cr.u@r
Loss:&s: inue:;;as;e as eteterior t&mperature decre.ises
-200 -1 00 1 00
Env1ronm e-ntal Temperature. c
National Aeronautics and Space Administration
As-is• Radiative coupling
• No additional layers
• Heat source temperature = 643 C
• Stirling hot-end temperature = 450 C
Minor modification• Compliant thermal interface
• Heat source temperature = 450 C
• Stirling hot-end temperature = 450 C
• Appears it could work for a 10-W DRPS with minor modification
Exploratory cases for 10-W DRPS
8
Thermal Simulator
• 643 450
600
400
500 350
300
400
250
300
200
200 150
100
100
so "' 50
Insulation Evolution
9
0.83 watts Hot RHU Surlace 500 C
Consistent Emissivities for 20 Year Life of Generator
Multilayer Insulation Overall Effective Thermal Conductivity-0.001 W/m-K
Sink Temperatures: Microporous Insulation Case A: -270 C (Space) Case B: -50 C (Cold Mars) Case C: +50 C (Hot Mars)
figure 1: MLMI in Cross Section
Coocept Only
Analysis needed for Deflnitizatkm
Cartridge
Heaters
Ef fective
Thermal
Conductivity
... 0.001
W / mK
Case B : -5()"( (Cold Mars)
Case C: •SO"C (Hot Mar5)
10-LAYERED PAIR ML! PEREGR!Nf COt.F•GURA r•O,\ '
RHU HOLDER '
l 0
400•c.-·e1 _.._ .. .....
Elcure 1; MLMI ln CrOH Section
-~--
National Aeronautics and Space Administration
Convertor ModelingConfidence in Predictions• 1D Sage vs. 3D CFD
• Modeled domain was truncated at the piston face and displacer rod (no seals, no bounce space, no displacer gas or radiation)
• Sage connects fixed temperatures directly to ends of the displacer, which artificially elevates displacer temps and associated axial parasitic heat transfer losses, while Fluent model resolves complex thermal and fluid flow fields
• Sage assumes no motion by the displacer when resolving heat flux while Fluent resolves temperature gradients and heat flux by moving components and deforming gas volume meshes
Codes agree well • The PV power output agree within 2%
• Indicated power predicted at 1.5 watts
10
Tcmpora1uro 625.0 609.2 500 4 577.6 581.8 548,1 530,3 514,5 498.7 482,9 487. 1 451 .3 435.5 419.7 403,9 388.2 372.4 3~.6 3'10.8 325.0
IKJ
875,000
850,000
825,000
_800,000 &. - 775,000 ~ ~ 750,000 ~
<>- 725,000
700,000
675,000
650,000 0.00
~ Piston Face
Gas Duct
Rejector Displacer Acceptor
I - PP -> Bl (Fluent) - CSD ·> BL (Fluent) - ESD ·> BL {Fluent} --- PP-> BL (Sage) --- CSD --> BL (S.igc) --- ESD -> BL {Sage)
0.10 0.20 0.30 0.40 0.50
Volume ( cm3 }
National Aeronautics and Space Administration
Objectives• Rectify AC power for load focused on controller design with generic load• Provide 1 We DC on 5 Vdc bus for rechargeable battery system and sensors• Charging battery enables house keeping power and periodic burst transmission of
measured data for communication
11
Controller Design
• Keep engine at a constant power level
• Shunt excess power when battery is fully charged and power is not required by the load
Stirlin,g converter
sense
AC to DC rectification
House keeping supply
Shunt
Battery
Transmitter
Load power regulation
RS422 to Controller vehicle or
data in GSE
Microprocessor
Load Sensors
c:::::J Cont roller hardware
c:::::J Load hardware
*optional ci rcuit if it proves to be energy efficient
National Aeronautics and Space Administration
Initial testing met the power output goals• Demonstrated linear AC regulator controller using a MOSFET H-bridge • Analog circuit controlling FETs for AC to DC rectification and alternator current control to improve
power factor • Load voltage monitoring allows for load control and shunting of unused power
12
Controller Breadboard Testing
Value Ideal diode rectifier
Wave form smoothing
Alternator Voltage, Vp-p
25.3 25.4
Alternator Power, We 1.37 1.34
Controller voltage, Vdc 11.7 11.5
Controller Power, We 1.11 1.22
Controller efficiency 80% 91%
CH3 5. OOV/d iv Position - 1.00 div
1: 111 / . ', t~I\
Output voltage
I
Estimated alternator
voltage
'v
Marn:10,111.
AcqModo 1 ; i Norr.tHl -500kS/s ?~s/rl Iv
[--l'ILE S<1ve
I • Filn I ist
-~ I
Dutil Type
-------------- ;..---··--=-""-=-- -~ ~~- -- ----- - -- ------- 4 rormdt
pp Avg
,R~,--, ' Slopped
Alternator voltage
l. t'.l~l°I
Cll1 0 .620A CH9 -' ,, .,. 6.51622V Cll:l ,, r, .,,.-. 7 .13393V
pp Avg Out y
.,-------
:Cll3 :CHI O :CII?
Alternator current
20 .18V 163.7 10nA 47 80X
Av11 :Cll4 RMS :CHI
3 .58%4nV 193 .14 3nA
Frame
l:xecute Save
3?8? -----------""' Fd~ Cll1"'.f"-------------•,!t:Fil•;;------2018/09/11 15: 30:23. 80529113 Auto 0.020A 2018/09/11 15:31 :01
National Aeronautics and Space Administration
Component Testing
13
Flexure Testing • FEA predicted static stress @deflection, corresponding stress noted for each
amplitude to develop S-N curve (fatigue stress vs. number of cycles to failure)
• Both flexure designs easily exceeded 10 million cycles operating at 100 Hz, high end of typical range used as threshold for identifying transition from high cycle fatigue to infinite life. The boundary between the finite-life and infinite-life regimes, or endurance limit, lies somewhere between 1 million and 10 million cycles for steels
• Displacer and piston flexures demonstrated over 100 million cycles with margin over the nominal amplitude without fracture
• Sufficient confidence our test effort will be failure free
Alternator Testing • Successfully demonstrated (2x) styles of 1-watt moving coil linear alternators,
(motored up to ~4 mm amplitude)
• Successfully demonstrated conduction path across flexure stacks and connections
• Demonstrated non-contacting operation of power piston
• Completed alternator design in concert with controller design in order to tune the motor inductance and ensure flight rated components (including capacitors)
Sandvik 7c27mo2 1095 spring steel
1400
1200
1000
i_ ,oo
j 600
400
200
HardUmll Amplitude:,3:m::;m;---------- I Nomlnal Amplitude, 2 mm
0 1.E+OO 1.E+02 1.E+06
1.E+0,4 des to F.ailure Number of Cy
1.E+08 1.E•10
National Aeronautics and Space Administration 14
Convertor DescriptionDesign Description
• Split-Stirling, gas duct between engine and alternator compression space
• Moving coil alternator
• Gap regenerator
• Flexure bearings for piston and displacer
• Laboratory design emphasized flexibility (not low mass)
Parameters• Hot-end temperature, 350 ºC
• Cold-end temperature, 0-50 ºC
• Operating frequency, 100 Hz
• Operating pressure, 110 psig
• Pressure amplitude, 13 psig
• Piston amplitude, 4.5 mm
• Displacer amplitude, 2 mm
Lab Test Setup(insulation not shown)
Alternator
Cartridge heater
Heat rejection loop
Gas duct
Engine heater head
National Aeronautics and Space Administration
Measurements• Hot and cold end thermocouples (8x)
• Dynamic CS pressure transducer (1x)
• Mean pressure transducer (1x)
• Hall effect sensors (2x)
• Power meter (power, voltage, current) for electrical heat input and alternator power output
Fabrication and Testing• Fabrication is complete
• Assembly is complete
• Pneumatic testing is complete
• Motor tests have started, some iterative changes are in progress
Convertor InstrumentationHot-end Temperature
Cold-end Temperature
Dynamic Pressure
Mean Pressure
Hall Effect Sensor
Hall Effect Sensor
Electric power input
Electric power output
15
-I -
I I
\ , I , -~
' I ' I ' 't,f !
' I ~
I I ' w - I , I
[ ~ ' ' I ' I - -
L _j . .
National Aeronautics and Space Administration 16
National Aeronautics and Space Administration
2019 NETS (Y. Lee, et al.)• JPL’s Small RPS A-Team Study focused on mission
concepts for (1 mWe to 40 We)
• Concepts were generated and reviewed as part of a JPL Innovation Foundry Architecture Team (A-Team) study
• Provided recommendations to RPS Program development activities
• Power, Mass, and Size (of interest for us)
• 1 to 10 watts
• 2.5 to 4 kg
• 3000 cm3 to 4000 cm3 (3U to 4U)
• Best fit for GRC 1-W DRPS1. Lunar Geophysical Network (10 We)
2. Pluto Lander (10 We)
3. Magnetosphere Study Fleet (1.5 We)
An Exploration of Mission Concepts That Could Utilize Small RPS
17
Mission Concepts Power Spectrum
Lmw. 1omw. 100mW. 10w. 40W.
Mission Concepts Mass Spectrum
0.5 kg 1kg 1.5 kg 2 kg 3 kg 3.5 kg 4 kg 4.5 kg
1000 cm• (•1u •) 6000 cm• ("6u• i 12000 cm:1r12U")
National Aeronautics and Space Administration
Conclusion• Small RPS concepts are being studied by NASA for use on small spacecraft
• GRC is working toward demonstration of a 1-watt dynamic RPS• Testing is in progress on all subassemblies (convertor, controller, insulation)
• Completed controller breadboard design and test, Brassboard is next
• Completed convertor assembly and alignment of moving components, 1st operation is imminent
• Completed insulation fabrication, test setup in progress
• GRC is also working on a paper design of 10-watt dynamic RPS
Small Stirling Technology Exploration Power or “smallSTEP”
18
----- - - -
National Aeronautics and Space Administration
Special thanks to:• RPS Program and Project Management• 1W Stirling Team
• Scott Wilson (Technical Lead)• Nick Schifer (Convertor Fabrication & Test)• Steve Geng (Alternator Design & Test)• Mike Casciani (Controller Design & Test)• Daniel Goodell (Thermal Management Design & Test)• Terry Reid (Advanced Modeling)• Barry Penswick (Engine Modeling & Design)• Paul Schmitz (Requirements)• Roy Tew (Sage Analysis)