i 

I 

I 

NASAlTM-20 13 -2165 61 AIAA-2013-3814 

Small Radioisotope Power System Testing 

at NASA Glenn Research Center 

Gina Dugala 

Glenn Research Center, Cleveland, Ohio 

Mark Be!! 

Vantage Partners, LLC, Brook Park, Ohio 

Salvatore Oriti 

Glenn Research Center, Cleveland, Ohio 

Martin Fraeman, David Frankford, and Dennis Duven 

The Johns Hopkins University, Applied Physics Laboratory, Laurel, Maryland 

November 2013 

l ___ _ _J 
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NASAlTM-20 13 -216561 AIAA- 2013- 3814 

Small Radioisotope Power System Testing 

at NASA Glenn Research Center 

Gina Dugala 

Glenn Research Center, Cleveland, Ohio 

Mark Be!! 

Vantage Partners, LLC, Brook Park, Ohio 

Salvatore Oriti 

Glenn Research Center, Cleveland, Ohio 

Martin Fraeman, David Frankford, and Dennis Duven 

The Johns Hopkins University, Applied Physics Laboratory, Laurel, Maryland 

Prepared for the 

11 th International Energy Conversion Engineering Conference (IECEC) 

sponsored by the American Institute of Aeronautics and Astronautics 

San Jose, California, July 14-17, 2013 

ational Aeronautics and 

Space Administration 

Gl enn Research Center 

Cleveland, Ohio 44135 

November 2013 
Acknowledgmen ts 

This work is funded through the NASA Science Mission Directorate. The authors wish to acknowledge personnel at 

NASA Glerm Research Center, The Johns Hopkins University Applied Physics Laboratory (JHU/APL), 

and Sunpower, Inc. , for their support and contributions to this project. 

Level of Review: This material has been technically reviewed by technical management. 

NASA Center for Aerospace Information 

7115 Standard Drive 

Hanover, MD 21076- 1320 

Available from 

National Technical Infonnation Service 

5301 Shawnee Road 

Alexandria, VA 22312 

Available electronically at http://www.sti.nasa.gov 
Small Radioisotope Power System Testing 

at NASA Glenn Research Center 

Gina Dugala 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Mark Bell 

Vantage Partners, LLC 

Brook Park, Ohio 44142 

Salvatore Oriti 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Martin Fraeman, David Frankford, and Dennis Duven 

The Johns Hopkins University 

Applied Physics Laboratory 

Laurel, Maryland 20723 

Abstract 

In April 2009, NASA Glenn Research Center fonned an integrated product team (IPT) to develop a 

Small Radioisotope Power System (SRPS) utilizing a single Advanced Stirling Convertor (ASC) with a 

passive balancer. A single ASC produces approximately 80 We making this system advantageous for 

small distributed lunar science stations. The IPT consists of Sunpower, Inc., to provide the single ASC 

with a passive balancer, The Johns Hopkins University Applied Physics Laboratory (JHU/APL) to design 

an engineering model Single Convertor Controller (SCC) for an ASC with a passive balancer, and NASA 

Glenn to provide technical support to these tasks and to develop a simulated lunar lander test stand. The 

single ASC with a passive balancer, simulated !tillar test stand, and SCC were delivered to Glenn and 

were tested as a system. The testing sequence at Glenn included SCC fault tolerance, integration, 

electromagnetic interference (EMI), vibration, and extended operation testing. The SCC fault tolerance 

test characterized the SCC's ability to handle various fault conditions, including high or low bus power 

consumptions, total open load or short circuit, and replacing a failed SCC card while the backup 

maintains control of the ASC. The integrated test characterized the behavior of the system across a range 

of operating conditions, including variations in cold-end temperature and piston amplitude, including the 

emitted vibration to both the sensors on the lunar lander and the lunar surface. The EMI test characterized 

the alternating cunent (AC) and direct cun-ent (DC) magnetic and electric fields emitted by the single 

ASC. The vibration test confin11S the SCC's ability to control the single ASC during launch. The 

extended operation test allows data to be collected over a period of thousands of hours to obtain long-term 

performance data of the ASC with a passive balancer and the SCc. This paper will discuss the results of 

each these tests. 

Nomenclature 

AC alternating cunent 

APS ASC piston sensor 

NASN TM- 2013-216561 

----- --- -
ASC 

ASC-E2 

ASC-L 

ASRG 

CSAF 

DAQ 

DoE 

DC 

EM 

EMF 

EMI 

FMEA 

FPC 

FPGA 

ILN 

IPS 

JHU/APL 

LED 

MEDA 

PID 

PV 

PWM 

RMS 

SCC 

SRL 

SRPS 

UPS 

We 

Advanced Stirling Convertor 

second-generation-Ievel Advanced Stirling Convertor 

Lunar-Derived Advanced Stirling Convertor 

Advanced Stirling Radioisotope Generator 

cold-side adapter flange 

data acquisition 

Department of Energy 

direct current 

engineering model 

electromotive force 

electromagnetic interference 

Failure Mode and Effects Analysis 

failsafe protection circuit 

field programmable gate array 

International Lunar Network 

integrated product team 

Johns Hopkins University/Applied Physics Laboratory 

light-emitting diode 

Macintyre Electronic Design Associates 

proportional integral derivative 

pressure vessel 

pulse width modulation 

root mean sq uare 

Single Convertor Controller 

Stirling Research Laboratory 

Small Radioisotope Power System 

uninterruptible power supply 

Watt electric 

1.0 Introduction 

NASA Glenn Research Center is the lead NASA center for the development of the Advanced Stirling 

Conveltor CASC) for use in space. In April 2009, NASA Glenn formed an integrated product team CIPT) 

to develop a Small Radioisotope Power System (SRPS) utilizing a single ASC with a passive balancer for 

possible use by the International Lunar etwork (ILN) program. At that time, the ILN program was 

studying the feasibility of implementing a multiple-node seismometer network to investigate the internal 

lunar structure. A single ASC produces approximately 80 We and could potentially supply sufficient 

power for that application. The IPT consists of Sunpower, Inc., to provide an ASC with a balancer, The 

Johns Hopkins University Applied Physics Laboratory (JHU/APL) to design an engineering model (EM) 

Single Convertor Controller (SCC) for the single ASC with a balancer, and NASA Glenn to provide 

technical SUppOlt to these tasks and to develop a simulated lunar lander test stand. 

The requirements for the SRPS differ from those driving the design and implementation of the 

Department of Energy (DoE) and NASA's Advanced Radioisotope Stirling Generator (ASRG). The main 

difference is that the required output power is modest and could be supplied by a single ASC-type Stirling 

convertor using half as much radioisotope fuel. Other differences from the ASRG include higher coldside adapter temperature and the need to control one rather than two convertor power sources. 

NASNTM- 2013-216561 2 
The Lunar-Derived Advanced Stirling Convertor (ASC-L) and SCC were developed to address these 

differences. The ASC alternator magnet was modified to operate at higher temperatures and a passive 

balancer was added to reduce vibration forces exerted by the convelior on the host. A fault-tolerant 

electronic controller, the SCC, for a single ASC was also developed. The controller included provisions to 

fully protect the convertor from single point failures in the controller and spacecraft. The controller also 

contains functionality to fully protect the spacecraft from single point failures in the convertor and 

controller. In either case, after a failure is corrected, the SCC design supports continuation of the mission. 

A lunar lander simulated test stand was developed to characterize the vibration produced by an ASC and 

understand its impact on a lunar lander. Several tests were performed on the system and individual 

components in order to reduce the risks associated with developing a small radioisotope power source for 

space (primarily lunar surface) applications based on using a single Stirling convertor. The results to these 

tests are presented in this paper. 

2.0 Small Radioisotope Power System (SRPS) Components 

2.1 Lunar-Derived Advanced Stirling Convertor (ASC-L) With Passive Balancer 

The ASC- E2 is a second generation of the engineering-level advanced Stirling convertor 

development by Sunpower, Inc. As a technology effort, the ASC-E2 design was evaluated to extend the 

current capabilities by investigating the adaptation of the technology to environments demanding higher 

rejection temperatures and higher alternator temperatures. Slmpower, Inc., built an ASC- L to demonstrate 

these capabilities. The ASC- L design is based on an ASC- E2 with limited changes and is intended for 

single-convertor configuration. The performance characteristics and temperature limits of the alternator 

were evaluated. 

The ASC- L differs from an ASC-E2 in four ways. First, higher temperature magnets were installed 

for an expanded cold-side adapter flange (CSAF) temperature range of 47 to 147 °C and a pressure vessel 

(PV) temperature range of 55 to 155 °C. In contrast, the ASC-E2s operated at a CSAF temperature range 

of 17 to 125 °C and a PV temperahrre range of21 to 132 °C. The rejection temperature range is based on 

simulations of expected lunar day/night temperatures. PV temperahrres were assumed to be 

approximately 8 °C higher than the rejection temperahrre based on analysis performed by Sunpower, Inc. 

The ASC- L will operate at high-rejection temperahlres for performance mapping only. Nominal 

operation is 850 °C hot-end and 57 °C rejection temperatures. Second, the epoxy used to bond the 

magnets is rated at a higher temperahrre than that on the ASC- E2. Third, the ASC-L includes a passive 

balancer (Figure 1) and does not require an interconnect tube or external mass for operation. Passive 

balancers are simple, under-damped, spring mass oscillators, harmonically hilled for sympathetic 

resonance. Fourth, the ASC- L is considered a developmental convertor since the same quality system 

used on the ASC- E2s was not implemented due to time and budget constraints. 

Limited design changes were made to the convertor to accommodate higher rejection temperahlre 

operation. The organics and materials for the internal limit sensor and piston were not changed for higher 

temperature operation but were deemed acceptable for short-tenn operation for 100 hr or less at a PV 

temperahlre of 155 °C. The ASC-L underwent various tests at NASA Glenn as described in the sections 

that follow. 

NASN TM- 2013-216561 3 

--~-~ - - --- -- -- --.---.-
I~~~-- Electric heater 

~--~-- ASC·L Housing 

Passive bala.ncer 

Figure 1.- ASC-L with passive balancer. 

2.2 Lunar Lander Simulated Test Stand 

The vibration produced by the ASC can affect the lander on which it is mounted. Pagnotta 

Engineering, Inc., was tasked with designing a test stand that simulated a lunar lander. This represents the 

first test stand in the Stirling Research Laboratory (SRL) at Glenn that attempts to characterize the effect 

of ASC vibration on a lander. 

To characterize the dynamic effect of a single, operating ASC on a landing platfonn, the test stand 

was designed to simulate the dynamics of a potential lander and measure the force at the interface 

between the ASC and the lander and at the interface between the lander and the planetary surface. 

Figure 2 depicts a conceptual design for an ILN lander. 

The test stand was designed to replicate the dynamic behavior of the conceptual lunar lander while 

avoiding the fundamental frequency of the ASC, 102.2 Hz, as well as the first two harmonics, with 

significant frequency separation, mimic the total mass of the lander of 150 kg, and provide isolation up to 

20 Hz for the potential seismometer instrumentation. Providing adjustable stiffness and damping at botb 

the interface between the ASC housing and the simulated lander and between the simulated lander and the 

planetary surface was also desirable to have the capability to simulate various landers. 

Figure 3 shows tbe final design for the lunar lander simulated test stand. It has a total estimated mass 

of 175 kg, a first vibrational mode at 30.36 Hz, a second mode at 320.2 Hz, and allows for adjustment of 

the stiffness and damping at both the interface between the ASC housing and lander simulator and 

between the lander simulator and ground. The test stand includes load cells at both of these interfaces for 

force measurement. While the ASC flexure utilizes a stainless steel plate in bending, the load fingers use 

a Sorbothane pad (Sorbo thane, Inc.) to provide both stiffness and damping. 

NASAlTM- 2013·216561 4 
NASA/TM-2013-216561 

--~---

1.0 m 

Notional 

Power ____ _ 

Svstem 

1 .... -----1.5 tYl -------.. 1 

Figure 2.- Conceptual design for ILN lander. 

,\'3[" t \\ 1111 PHI' 

II 1mA' t til 

M..clwtAtDl fl.ltJw.A/.., 

Figure 3.-Final design for lunar lander simulated test stand. 

5 
2.3 Single Convertor Controller 

The goal of the see task was to design and build a small, efficient, and reliable controller (Ref. 1) 

for use with a single ASe. The see maintains stable operation of the Stirling convertor. It regulates the 

alternating current CAC) produced by the linear alternator of the convertor, provides a specified direct 

current (DC) output voltage for the spacecraft, and maintains operation with a stable piston amplitude and 

hot-end temperature. The controller allows adjustment of the operating point; the hot-end temperature or 

piston amplitude may be increased or decreased. The piston amplitude is maintained by varying the load 

with the regulation stage of the controller. 

The see is designed for use with one ASe whereas the others used in the SRL at Glenn are intended 

for dual-opposed configured ASes. The controller for a dual-opposed pair of convertors must synchronize 

the piston motion of the two convertors to minimize vibration as well as manage their operation. The 

controller was designed for flight use, but only a mechanically equivalent, nonflight EM was built. 

Unlike most ASe controllers in the SRL, the see does not require a physical tuning capacitor to 

keep the stator current in phase with back electromotive force (EMF) voltage and piston velocity. These 

capacitors have a large mass and volwne and are unattractive for spacecraft applications. 

A block diagram of the see connected to a convertor and a typical redl.mdant spacecraft is shown in 

Figure 5. The see consists of two identical controller boards packaged in separate chassis. One tmit is 

actively in control of the convertor at all times, while the other is not powered lmless needed to recover 

from a fault. Each controller board contains the power handling, data acquisition, signal processing, and 

secondary voltage conditioning circuits needed to control a convertor and deliver De power to the 

spacecraft. 

The see not only provides power to the spacecraft but it also must regulate ASe operation to avoid 

damage to internal components and maintain safe thermal conditions after fueling. The controller was 

designed with full redundancy to avoid potential overheating or damage to the internal structure. During 

see development, a fl.illctional failure mode and effects analysis (FMEA) listed possible functional 

failures in the circuits. The impact of such failures and mechanisms to detect and correct them were 

identified and implemented in the detailed design. A fault-tolerant see consists of two identical boxes, 

each containing a copy of the same circuit. In typical operation, one box is actively controlling the 

convertor, while the other is available as a backup in the event of a fault. The current best estimate mass 

of both boxes, without the interconnecting harness, is 2.27 kg using an AI-6061 -T651 chassis and 2.00 kg 

ifMg-ZK60A-T5 material is used. Pictures of the see and EM circuit card in an assembly fixture are 

shown in Figure 4. The see underwent various tests at NASA Glenn as described in the sections that 

follow. 

Figure 4.- SCC and EM Circuit Board. 

NASN TM- 2013-21656 1 6 
SmRPS sec Spaeecraft 

. I P<m&r 

Y .. l' Power 

SIlum ~?CYI ~ Systems 

ShI..u\t ~ Elec trOl'lics Ras,~ r CoolJOller 

.... T~~Poo 

AC I CMV(fUl 

Command & 

.. I I:: .. ... . . "- I 

Data Handling 

f- .... TC$fPoo "-: A 

~ TP'I/ ~ ~~1 ' 

Coril!Ol1oer Res!$lJ)I' 

= Command & TAll PoYt~( ... Oatatfandlrng 

S 

Figure 5.-Block diagram of a typical see redundant spacecraft application. 

3.0 Lunar-Derived Advanced Stirling Convertor (ASC-L) Testing 

3.1 Temperature Performance Mapping 

Before delivering the ASC-L convertor to Glenn, Sunpower, Inc., completed the temperature 

performance map shown in Table 1. The perfonnance map test points and results produced at Sunpower, 

Inc., were replicated in June 2012 at Glenn as shown in Table 2. 

TABLE I - SUNPOWER INC ., ASC- L TEMPERATURE PERFORMANCE MAP DATA 

Mission phase Hot-end CSAF PV Alternator Alternator Alternator Piston 

temperature tempera ture temperature voltage, current, power, amplitude, 

+0 /-2 °C ±2 °C ±2 °C V A W mm 

Beginning of 847.2 89.5 97.9 14.14 10.37 75.2 4.09 

mission, max. 

Low reject 846.8 46.2 54.8 11.1 8 8.91 60.2 3.58 

High rej ect 846.9 146.0 152.3 16.62 11.85 68.0 3.99 

TABLE 2 - GLENN ASC- L TEMPERATURE PERFORMANCE MAP DATA 

Mis ion phase Hot-end CSAF PV Altern ato r Alternator Alternator Piston 

temperature temperature temperature vo ltage, current, power, amplitude, 

+0 / - 2 °C ±2 °C ±2°C V A W mm 

Beginning of 847.2 88.9 97.9 14.3 3 10.55 78.75 4.08 

mission, max. 

Low reject 846.5 46.9 54.1 11.36 9. 16 65.49 3.58 

High reject 846.4 144.2 152.9 16. 17 11.75 65.53 3.99 

The ASC- L was delivered to Glenn with the Sunpower, Inc., mechanical hardware design, which 

included an insulation package, heater assembly, CSAF cooling jacket, and housing. This hardware was 

designed to allow for higher rejection temperature operation up to 147 °C. Due to the limited ASC-L 

design changes for high-temperature operation, high-rejection temperature operation was only performed 

for the temperature perfonnance map; nominal convertor rejection temperature operation is 57 °C. The 

ASC- L was controlled by an AC bus power supply instead of the SCC since this type of controller was 

used by Sunpower, Inc. An AC bus power supply simulates the convertor being connected to an AC grid. 

The AC bus controller controls piston amplitude by varying the AC voltage and therefore, the load on the 

convertors. Increasing the AC voltage increases the piston amplitudes. The test results between Glenn and 

N ASAJTM- 20 13-21 656 1 7 

l 
Sunpower, Inc., are comparable with the exception of the alternator power. The difference in alternator 

power measured at Glenn and Sunpower, Inc., is due to the location of the alternator power measurement 

at the feedthrough versus the connector. Upon completion of tbis test, the ASC- L was prepared for other 

component-level and integrated tests as will be discussed in the sections that follow. 

3.2 Electromagnetic Interference (EM!) Testing 

The ASC-L was tested in February 2013 for radiated EMI while controlled by first an AC bus power 

supply and afterwards the SCC. The goal was to compare for a baseline against second-generation-Ievel 

Advanced Stirling Convertor (ASC-E2) radiated EMI test results and also to characterize the difference 

when controlled by the SCc. The baseline testing determined that there was, on average, an overall slight 

increase in AC magnetic field strength (as compared to the ASC-E2), but was otherwise very 

comparable. The EMI testing procedures were based on MIL- STD-461-C, referencing sections REOI, 

RE02, CSO 1, CS02, CS06, CEO 1, and CE03. AC magnetic field measurements were taken using a 

Macintyre Electronic Design Associates (MEDA) MGCH-2 magnetic probe with a MEDA SAM- 3 

amplifier as well as a 13.7-cm SAS-200/56 (450) loop antenna by AH Systems. DC magnetic field 

measurements were taken with a F.W. Bell HTM99-0608 Hall Effect Probe with a F.W. Bell 9550 Gauss 

Meter. AC electric field measurements were taken using an active rod antenna (EMCO 330lB) or a diskcone antenna (Antenna Research SAS-2Nm) depending on the frequency range being focused on. 

Conducted emission tests used AH Systems current probes, namely BCP-200/51 0 for low frequencies 

and BCP-200/511 for high frequencies. The conducted susceptibility testing used a Marconi 2031 

generator. 

During AC magnetic field testing, the SCC increased the radiated emissions from the ASC-L 

marginally 0.5 dBpT), but the main effect was an introduced spike at 23 kHz, the pulse width 

modulation (PWM) switching frequency of the SCC, plus harmonics. The data is contained in the NASA 

Glenn EMI Test Data Report GRC-EMI-RPT - 342 (this also contains the data for the SCC susceptibility 

and conducted emission tests mentioned later in this report). Figure 6 shows the difference in measured 

magnetic fields when ASC-L is controlled by the AC Power Supply as compared to the SCC. 

During AC electric field testing, the SCC decreased the 102.2 Hz dominant spike but increased the 

204.4 Hz spike, with associated harmonics. It also injected a 23-kHz spike (likely due to the PWM 

switching as mentioned above), and introduced broadband noise from 400 kHz through ~ 15 MHz. As 

this was a general characterization test, there has not been much effort to identify the sources of the 

changes introduced. Figure 7 and Figure 8 show the effect on axial electric field measurements between 

AC bus control and SCC control. 

l.o>-elldBpT) 

.00 

80 

40 

50 100 

Frequency (Hz] 

..................... ......................... -... . ......... . 

~ 1 ! l j [ 1 i j ~ 1 ; i ! ill; [ ; 

'iIIII]II,i{illll!!1 . -: .. :- { .. ~ 

::::l 

Flequency [Hz] 

Figure 5.-Axial AC magnetic fields, AC bus (left) versus SCC (right). 1 m (pink), 2 m (blue), and 3 m (green). 

NASN TM- 20 13-216561 8 
Frequency (HzI 

Figure 7.-Axial electric field measurement, AC bus control. 1 m (pink), 2 m (blue), and 3 m (green). 

SO 100 200 400 1~ 2k 4k 10k 10k lOOk 

Frequency (Hz] 

Figure 8.-Axial electric field measurement, SCC control. 1 m (pink), 2 m (blue), and 3 m (green). 

4.0 Single Convertor Control Testing 

4.1 Fault Testing 

The see not only provides power to the spacecraft but it also must regulate ASe operation to avoid 

damage to internal components and maintain safe thermal conditions after fueling . The controller was 

designed with full redundancy to avoid potential overheating or damage to the internal structure. 

Spacecraft power needs may briefly exceed the capacity of the SRPS power source . There are many 

causes of such events; examples include activating thruster valves, sudden momentum wheel speed 

changes, telemetry downlink transmission, in-rush cunent surges when activating loads, and bursts of 

science data collection are among many such possibilities. If auxiliary power is not supplied from another 

source, the power bus voltage may be pulled too low and trigger a spacecraft fault. A battery frequently 

acts as the auxiliary power source to ensure that an adequate bus voltage is maintained during these power 

transients. Alternatively, a large capacitor may provide backup power if the expected power transients are 

NASAlTM-20 13-216561 9 
short and the energy storage capacity of the capacitor is adequate. Thus, the SCC was required to support 

either a battery or capacitive power bus spacecraft architecture. 

During SCC development, a functional FMEA listed possible functional failures in the circuits. The 

impact of such failures and mechanisms to detect and correct them were identified and implemented in 

the detailed design. The scope of potential issues were not limited to internal controller failures but 

extended outside to include impacts from the spacecraft and the engine. Recovery from a fault is based on 

switching from the controller board that detected the fault to the backup board. After detecting the Hbridge fault and isolating the failed controller, modeling showed that the backup controller had less than 

20 ms to reestablish proper operation of the convertor. Modeling also showed that the open/short must be 

removed in less than 5 ms to avoid internal damage to convertor components. 

The fault detection and correction mechanisms in the controller protect the convertor and spacecraft 

from damage while recovering from a fault. Power flows from the convertor through an electronic switch 

that acts as an AC circuit breaker. The AC circuit breaker is opened in the event of a fault to isolate the 

convertor from the failure and allow the redundant board to initiate control. An electronic DC circuit 

breaker utilizing two FET switches will isolate the controller board from a spacecraft fault and isolate the 

spacecraft from an internal controller fault. While the DC breaker is open, convertor power flows to an 

external shunt resistor controlled using PWM to maintain a fixed H-bridge output voltage. A combination 

of analog circuits, field programmable gate array (FPGA) logic, and microcontTOller software monitor 

SCC operation and detect and correct potential faults. 

Analysis was completed to determine methods for achieving hot swap functionality. Hot swap refers 

to changing a failed controller card while the active controller card maintains operation of a convertor 

operating at full power. Due to hot swap capability, each controller card requires its own set of 

harnessing. 

The SCC successfully handled various fault conditions, including high or low spacecraft power 

consumption, total open load or short circuit, and replacing a failed SCC card while the backup maintains 

control of the ASC. These tests were completed with an ASC-1 #4, early Sunpower, Inc., developmental 

convertor. The main difference, as far as the controller is concerned, between the ASC-L and ASC-1 #4 

is the operating frequency. The ASC- L operates at a frequency of 102.2 Hz and ASC- 1 #4 at 103 Hz. 

However, the SCC can be commanded to operate in the frequency range of99.0 to 104.7 Hz. 

The SCC fault testing, in November 2011 and March 2012, included spacecraft disconnect, spacecraft 

bus voltage variation, AC circuit breaker open load, and hot swap. The results of these tests can be seen in 

Figure 9 to Figure 12. The spacecraft disconnect test confinned the SCC's ability to operate in the fueled 

configuration. During fueling, the SCC will be disconnected from the spacecraft and will dissipate power 

in a shlmt resistor. The DC bus voltage variation test confirmed the SCC's ability to maintain convertor 

power output stability over a range of DC bus voltages. The AC circuit breaker open load test is one of 

many open/short circuit fault tests perfonned on the SCC. This verifies the SCC's ability to handle an 

open/short either on the ASC or spacecraft side. The hot swap test confirn1ed the SCC's ability to 

maintain ASC control while replacing a failed controller card. Additional testing included verifying the 

ability of the SCC to maintain convertor control while going over and under the acceptable DC bus 

voltage range, successful power transfer from the spacecraft to the shunt resistor, and recovery from the 

fault. The SCC attempts to reconnect to the spacecraft and connects when the spacecraft voltage reaches 

an acceptable range. 

NASAlTM- 201 3-216561 10 

L_ 
OO r-------------------------------------------~~------~ 20 

~~--------------~-------+~-

M+-______________________ SIP ia ce ~r~ ft~ _____ \Shunt 

70 +_----------------------------------------~~------~ 

60 +-------------------------------------------~ ------_1 

3: 

50 +-------------------------------------------~ --------1 o 

0.. 

... 

o 

~40 +------------------------------------------- -~-------I Q) 

~ --------------------------------------~~--- 30 +_----------------------------------------~~-------1 

10 ·r--3£ -----------------------------~~

o f---------,-.----,- I 0 

14:52 14:55 14:58 15:01 15:04 15:07 15:10 15:12 15:15 15:18 1521 15:24 

TIme (hh:mm) 

Figure g.-Spacecraft bus voltage variation. Piston amplitude (blue), alternator 

voltage (red), alternator current (green), and alternator power (black). 

90 ,------------------------------------------------------------,- 20 

89 ~----- --------.. - .----------.. -----------.------.-.---------- . 18 ~ r ~ 

M +-----------.----~ 

a. 

E 16 « 

c 

{l 

87 +------==~====~==~~~~~~~==~~~~~~~~~ ------~ 14 ~ !ii Shunt « i 86 ------Spacecraff---- "---j 12 I ~ 

~ 85 10 ': ~ o o - M m 

E ~., .e 84 - .-.--. . .----------------------.-----.--- 8 _ 

~ ~-~ ~ 

83 t-----------------+-....... --------------------------------------t 6 ., 

01 

2 

o 

82 4-----========:::::==+==t:===========================------/- 4 :;: 

81 +--.-.... -.. -----.. -.. ---.----~=~~

, .J -.----.--------------------------------4 2 ~ ~ 

80 -----~---.,------~-- --~ __ --~ ---~---___,----_,. -- ·_· 0 

11 :06 11 :07 11:07 11 :07 11 :07 11:08 11 :08 11 :08 11 :09 11 :09 

TIme (hh :mm) 

Figure 1 D.-Disconnect from the spacecraft. Piston amplitude (blue), alternator voltage 

(red), alternator current (green), and alternator power (black). 

NASAlTM--20 13-216561 11 
80 .-.---------.---.------------.--------- --1 20 

E 79 +---------------------------------+ 18 ~ 

.. 

"0 

B 

78 +-- ---------- - ------------------+1 16 ~ 

I ..: 

~ :: +----------.----p E·-M- 3-------------t

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12:33 12:33 12:34 12:34 12:34 12:34 12:34 12:34 12 :35 12:35 

Time (hh :mm) 

Figure 11.- AC circuit breaker open load fault. Piston amplitude (blue). alternator voltage (red). 

alternator current (green). and alternator power (black). 

~ y-------------------------------------------------------, 

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14:24 14:25 14:26 14:26 14:27 14:28 

Time (hh:mm) 

Figure 12.-Hot swap. Piston amplitude (blue). alternator voltage (red). alternator current 

(green). and alternator power (black). 

NASAJTM-20 13-216561 12 
N 

I 

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Ol 

1.0000 

0.1000 

0.0100 

0.0010 

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, 

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Frequency, Hz 

Figure 13.- Random vibration profile. 

4.2 Flight Acceptance Vibration Testing 

10000.00 

The ASC-L was subj ected to flight-acceptance-Ievel vibration testing (Figure 12) under sce control 

in January 2013. The goal of the testing was to achieve qualification-level vibration. Flight-level vibration 

is -3 dB from qualification-level vibration. The convertor performed acceptably dming the flight-level 

vibration testing with no change in performance noted when compared to the pre-vibration test run. The 

see maintained safe piston margins when subjected to flight-acceptance levels. However, the postvibration data analysis noted that the ASC-L was producing very high levels of cmrent (near 25 Arms) 

during peak accelerations. This high level of current was near the saturation point of the see's current 

sensor, and is something that will be addressed in future controllers. Due to the high levels of current, the 

test ended at flight-acceptance level. This data was captured in the Glenn Structural Dynamics Lab Test 

Report SDL-TR-13-03. Figure 13 shows the random vibration profile that was used during the test. 

Due to the nonlinear output of the ASe piston sensor (APS), the real-time decision to abort dming a 

vibration test has relied in the past on either an altemate position sensor being used or on ancillary data 

(current, etc.). The data acquisition (DAQ) system in the test rack samples during a 2-sec window and 

outputs a calculated value-clearly not appropriate for a real-time abort decision on a random vibration 

test. An independent nearly real-time APS processor was developed to provide a more linear piston 

position output that could be viewed on an oscilloscope to make abort decisions during the test run based 

purely on piston proximity to bard stops inside the ASC. Analysis of a data post test sbowed that while 

the processor output lagged an actual piston peak position by 3 to 4 cycles (at 102.2 Hz), the data was 

provided to the operator quickly enough to shut down the test before any actual contact events occurred. 

NASNTM- 2013-216561 13 
4.3 Susceptibility and Conducted Emissions Testing 

The SCC was tested for conducted line emissions while operating the ASC-L using test procedures 

based on MIL-SPEC-461C CE01 and CE03. The majority of peaks were detected at -204 Hz (two times 

convertor operating frequency) with 4 to 5 haIIDonics, at - 23 kHz (approx. SCC PWM switching 

frequency) with several harmonics and at - 700 kHz. There was no discernible difference in conducted 

emissions between the different operating modes of the SCC nor on the power versus return leads. There 

was a significantly higher (-40 dB A) level of noise pre-capacitor bank versus post-capacitor bank as 

shown in Figure 14. 

Additionally, the SCC was subjected to conducted EMI susceptibility testing using test procedures 

based on MIL-STD- 46IC CSOI, CS02, and CS06. To protect the ASC- L hardware against a loss of 

control event, the convertor simulator was used to provide power to the SCC and a resistive test load was 

used to pull full power during the duration of the test without using an electronic load. The convertor 

simulator was developed by APL and simulates the operation and electrical behavior of an ASC in real 

time via a combination of hardware and software. The CSO 1 and CS02 tests are power limited to 1 Wand 

voltage limited to 1 Vrms-both tests typically experienced voltage limiting in the lower kHz range during 

CS02 where power limiting dominated. During conducted susceptibility testing, there were no noticeable 

performance effects. Output voltage change, controller latch-up, or loss of communication were all 

monitored and did not occur. Figure 15 details the profile of two of the pulses applied to the power path 

during CS06 and are similar to other pulse events. 

Leve!rd8~) 

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0 20 40 60 100 200 400 1k 2k 3k 5k 20k 

Frequency [Hz) 

0~10~k~W~k~~100~k~WO~k~~1~M~2M~~4~M~~0~~~~2~O~~J~OO~M 

Frequency (Hz] 

- MES CE01f - MES CE03f 

-MES CEOlg -MES CE03h 

Figure 14.- CE01 and CE03 comparison of pre-capacitor (blue) and post-capacitor banks (green). 

NASA/TM- 2013-216561 14 
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Figure 15.-eS06 positive and negative pulse on power return lead. 

NASAlTM- 20 J 3-2 J 6561 15 
5.0 Integrated Testing 

5.1 Passive Balancer Tuning 

The ASCs tested in the SRL at NASA Glenn are tested both as single units and dual-opposed pairs. A 

single ASC requires mounting to a large mass to reduce vibration produced by the convertor. A dualopposed pair is mounted with the heater heads facing outward and the PV sections rigidly attached to 

each other. This configuration permits dynamically balanced operation where the piston's motions are 

equal but opposite in direction. However, a single ASC mounted to a large mass is impractical for 

spacecraft applications. Therefore, under the SRPS project, Sunpower, Inc., designed a passive balancer 

to replace the large mass and spring system used to absorb the vibration produced by the ASC. This is the 

first integration of an ASC with a balancer. 

The passive balancer is tuned to balance the convertor force at a specific operating point, including 

piston amplitude, rejection temperature, balancer temperature, and convertor operating frequency. The 

ASC-L with the passive balancer is mounted to the llmar lander simulated test stand. The load cells 

mounted at the base of the test stand and between the simulated lunar lander simulated test stand and the 

ASC-L housing provide data on the vibration emitted by the convertor to the lander and lunar surface. 

Prior to tuning the passive balancer in May 2103, a tap test was performed at Glenn to characterize 

the lunar lander simulated test stand. During this test, several accelerometers mounted on the various 

components of the test stand, including the ASC-L housing, the lunar lander mass simulator, and the 

floor, recorded the accelerations resulting from taps imparted to the structure at different locations. This 

test verified that the frequencies and shapes of the system modes match those predicted in the design 

analysis. 

When designing the passive balancer, the dynamics of the lunar lander simulated test stand must also 

be considered. The operation of the balancer is effected by the compliance between the ASC and the test 

stand. If the compliance is greater than 9000 N/mm then the balancer will not work as a vibration 

canceller. If the compliance is less than 9000 Nlmm, the vibration characteristic of both the test stand and 

balancer is varied by the balancer. At the time the test stand was designed this compliance requirement 

was not known and therefore was designed with compliance greater than 9000 N/mm. Since, as 

mentioned in Section 2.0, the lunar lander test stand was designed with an adjustable stiffness this 

compliance could be changed. After this adjustment, the passive balancer was tuned to operate at 850 °C 

hot-end temperahlre, 57 °C rejection temperature, 4.25-mm piston amplitude, and 102.2 Hz operating 

frequency. This operating point resulted in a passive balancer temperature of35 DC. As part of the tuning 

process, an accelerometer mounted on the ASC-L housing was used to find an operating frequency that 

minimized acceleration. The convertor operating frequency was changed, while the convertor was stalled, 

in 0.01 Hz increments up to ±0.5 Hz deviation from the convertor operating frequency. Once the 

frequency was detemlined, a calculation was performed to determine the mass needed to minimize the 

force produced by the ASC. This mass was installed on the ASC to characterize its ability to reduce the 

force exerted on the lunar lander test stand and planetary surface. 

The tuning resulted in a force reduction between the ASC-L housing and lunar lander simulated test 

stand from 253.4 to 111 .6 Newtons and a reduction between the llmar lander and planetary surface from 

265.8 to 118.7 Newtons. Options to further reduce this force are being considered. The SRPS operates 

24-br17 days per week without the passive balancer since the temperature regulation in the lab is not 

suitable to meet the balancer temperature limitation of ±4 DC. If the balancer temperature limitation is 

exceeded, the balancer will resonate and cause potential damage to the ASC-L. Other balancer options 

without this temperature limitation are being considered. 

NASA/TM-2013-216561 16 
--- Test Rack 

o1----ASC-L 

Wateriglycol bath 

Lunar LlIllder 

Simulated Test Stand 

Figure 16.-SRPS extended operation setup. 

5.2 Extended Operation 

Extended operation (Ref. 2) of ASCs has been an ongoing effort at NASA Glenn since 2003. This 

effort was initiated to demonstrate the life and reliability of Stirling convertors. Extended operation 

involves continuous, unattended convertor operation over a period of thousands of hours. The test rack 

provides the means for measuring and collecting data for observing operating trends. In addition, it 

provides convertor control and safe operation. Figure 16 shows the SRPS test rack in Glenn's SRL. 

The SRPS test rack is designed to operate either an AC bus power supply controller or the SCc. Each 

test rack allows the adjustment of operating conditions, including hot-end temperature, rejection temperature, and piston amplitude. Each test article was instrumented to measure these parameters in addition to 

output characteristics such as alternator voltage, current, power, and operating frequency. The hot end of 

the convertor is heated with electric cartridge heaters (Figure 1), and the cold end is maintained with a 

water/glycol bath. 

The data system utilizes National Instruments' Lab VIEW -based DAQ hardware and software to 

acquire data and monitor the test. It displays and records data on a computer, collects and saves data in 

various timeframes, calculates parameters with received data, and provides safety to the convertors. The 

DAQ software, which was developed to operate in unattended mode, can control the support systems 

without user intervention. 

Data collection allows for detailed analysis of convertor operation. To achieve this, LabVIEW 

collects and displays several parameters including 

• Heater voltage and cutTent 

• Electric heat-source temperature 

• Hot-end temperature 

• Cold-end temperature 

• Ambient temperature 

• PV temperature 

• Alternator root mean square (RMS) voltage, RMS cUlTent, and power 

• Piston amplitude 

• Operating frequency 

NASN TM- 2013-216561 17 
• AC bus power supply controller voltage and current 

• Power factor 

• Heater power 

A complete data record of these parameters is stored every hour. In addition, the DAQ system 

maintains a buffer of data recorded over the last 24 hr at a 2-sec scan rate, records an average of the data 

over the last 5 min of every hour, and allows these data to be stored manually as needed for detailed 

analysis. From the stored data, numerous parameters are calculated including 

• Heater resistance and power 

• System efficiency 

• A verage hot-end temperature 

• Cold-end temperature 

• Piston amplitude 

• Net heat input 

The DAQ software collects and displays various telemetry parameters. In addition, it allows for the 

following SCC commanding: 

• Changing the voltage applied to the convertor to start piston motion, Vm 

• Convertor operating frequency 

• Disconnect from the spacecraft 

• Connect to the spacecraft 

• Operation with or without the backup controller card 

• Open load or short circuit fault injection 

Convertor hot-end temperature is controlled by a hot-end thermocouple input to a Partlow 

proportional integral derivative (PID) temperature controller. The PID temperature controller varies the 

voltage of a progranunable Xantrex DC power supply connected to cartridge heaters in a heat source 

block to maintain constant hot-end temperature. The PID temperature controller is programmed to limit 

the rate of change in temperature to 1 °C/sec. 

Electronics in the SCC place requirements on the power path line resistance. The test rack 

accommodates this requirement in two ways. First, the power panel is mounted as close to the convertor 

as possible to reduce the length of wire in the power flow path. Second, external current sensors are used 

for current measmements. These modifications allow measmements to be taken close to the convertor. 

The SCC relies on a DC electronic load to sink power in an output voltage range between 22.0 and 

36.0 V de and is determined by the end user. The SCC will track to this voltage providing power as a 

constant-power source, and will sense a bus overvoltage condition. The SCC will provide constant power 

as long as the end user is in the required voltage range, but it will intemally shunt power and disconnect 

itself from the bus if the voltage range is exceeded. Shunt resistors sized for maximum output power of 

the convertors are mounted on the test stand for this purpose. The bus is connected to a capacitor bank: for 

transient energy storage that allows a range of 10 000 to 100 000 flF. 

Safety systems on the test racks are implemented through the DAQ system and through hard-wired 

circuitry. The user may specify upper and lower bounds for any parameter monitored by the DAQ system. 

The DAQ software will safely shut down operation of the test article when an out-of-bound condition is 

sensed. Parameters that trigger the shutdown of the ASC-L include low or high hot-end temperatures, low 

or high rej ection temperatures, bigh PV temperature, bigh piston amplitude, and loss of building power. If 

a shutdown is initiated, the DAQ software commands the hot-end temperature to 10 °C and the circulator 

temperature to 25 DC, and it turns off the heater power supplies. The convertors continue to be motored 

with the SCC until the average hot-end temperature reaches 700 °C, with the nominal hot-end 

NASN TM- 20 13-216561 18 
temperature being 850 °C. A shutdown is initiated for loss of building power after 5 min of operating on 

an uninterruptible power supply (UPS) system. Another safety feature in the software notifies the end user 

when a shutdown occurs, logs the fault via the DAQ, and records the data. The end user is notified by a 

cell phone text message that indicates the fault code of the shutdown. 

Two hard-wired protection devices installed in the operations rack function independently of the 

software-based protection: temperature limit and failsafe protection circuit (FPC). If a fault is registered 

by the temperature limit device, the DAQ, or an optional third source, the heater power supplies are 

disabled and a fault is indicated by a light-emitting diode (LED). If the hot-end temperature of either 

convertor exceeds the user-defined limit, the limit controller removes the heater power from both 

convertors via a relay. A FPC was implemented to prevent piston overstroke. The FPC can monitor up to 

four input signals. Each input has an associated, user-adjustable setpoint. When any signal exceeds its 

setpoint, the convertor controller is disconnected, and an emergency load is applied across the alternator 

and a signal is sent to the DAQ software. A protection circuit interface panel provides a means for the 

user to be notified via LEDs when a piston amplitude limit is exceeded. The user can reset the protection 

circuit after the fault and can manually stall the convertors by flipping a switch on this panel. The user 

also can disconnect the protection circuit and select between AC bus power supply control and the SCc. 

The SRPS has been in extended operation since May 20l3. To date, no noticeable changes have been 

identified in system performance. The system performance data, as shown in Figure 17 and Figure 18, is 

changing with the ambient temperature in the SRL. The temperature regulation in the laboratory is not 

controlled as desired so ambient temperature changes are common. The figures also indicate a loss in data 

collection for approximately 1 day due to a computer malfunction. The system performance data will 

continue to be analyzed for thousands of hours. Any noticeable changes in performance will be identified 

and if possible corrected. 

850 .................. 1' ............... .. 75 

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Figure 17.-ASC-L temperature performance data. 

NASN TM-2013-21 6561 19 

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00 50 100 150 200 250 300 350 4004 

Hours of Operation 

Figure 1B.-ASC-L alternator and piston amplitude performance data. 

6.0 Discussion 

The testing described in this paper sought to reduce the risks associated with developing a small 

radioisotope power source for space (primarily lunar surface) applications based on using a single Stirling 

convertor. 

Magnetic and electric field measurements of the ASC-L were similar to the baseline ASC-E2. Field 

levels were not significantly different when the ASC-L was controlled with the SCC or a simple AC-bustype controller. 

Fault tolerance of the SCC was extensively exercised. A wide range of spacecraft and ASC faults 

were introduced while operating at full power level. In addition, many different faults were also injected 

internal to the controller. In all cases, the fault was detected, a corrective action taken, and telemetry 

updated with diagnostic information on the event and response. ASC piston amplitude was controlled 

during all these fault cases. While piston amplitude had transient changes during some fault responses, 

careful design of the fault detection and response algorithms meant the piston never hit the hard stop even 

when switching controller boards. 

External disturbances such as vibration during launch affect motion of the ASC case. The alternator 

magnet is directly connected to the convertor piston while the alternator coil is paIi of the external case. 

Hence, external disturbances result in ASC current output disturbances as the coil is moved relative to the 

magnet. Vibration testing at flight levels showed that current variations approached the limits of the 

controller's full digitization range. Exceeding that range would cause the control algorithm to process 

inaccurate input current values and therefore, risk loss of control. To allow higher external vibration 

levels, the SCC input alternator cun-ent dynamic range should be increased. However, if dynamic range is 

increased, resolution will be decreased and ASC output power level noise may degrade. Hence, an 

analysis to predict maximum input current while under external vibration should also be performed to 

support optimizing the dynamic range/noise tradeoff. 

Conducted ernissions testing of the SCC showed a large current component at twice the ASC 

operating frequency. The current is caused by the controller's full wave rectification of the ASC output 

power. For a capacitive bus-based electrical power system, the bus capacitance would smooth the current 

and minimize ernis ions. A battery-based power system would also reduce emissions but in that case the 

NASN TM- 20 13-216561 20 
ripple current would also be equivalent to charge/discharge cycles at a high frequency. The impact on 

battery life of such treatment is unknown so design work to add an additional output regulator stage to 

remove most ripple current has started. Simulations of that preliminary design show that it should be 

possible to lower output current ripple below a few lOs of rnA at the expense of additional power loss in 

the controller. 

Open loop testing of the ASC-L passive balancer showed a force reduction between the convertor and 

spacecraft of somewhat more than a factor of 2. However, the passive balancer is only helpful over a 

fairly narrow ambient temperature range. Conceptually, a control algorithm may be possible that senses 

force (or related parameter) between the convertor and then changes operating conditions, such as piston 

oscillation frequency, to maintain the effectiveness of the passive balancer over a wider temperature 

range. The ASRG reduces force far more than the passive balancer by using a second ASC with all 

internal masses moving in the opposite direction to the first ASC so the net momentum across both 

convertors is cancelled to the extent that ASC characteristics are matched. Conceptually, for a single 

ASC-based power source, an actively driven balance mass could be used to similarly cancel convertor 

momentum. The active balancer would require power to drive its balance mass as well as control 

electronics to match ASC motion. Hence, less power will be available for the load but that may be an 

acceptable trade for some applications. 

The test results we have presented here show that is feasible to develop an SRPS suitable for a small 

lunar-Iander- based seismometer network. The largest remaining challenge is developing a mechanical 

interface between the SRPS and lander that does not allow vibration from the SRPS to affect the quality 

of collected science data. Controller improvements to reduce current ripple, tolerate launch vibration, and 

balancer effectiveness would also be helpful. 

7.0 Conclusion 

The Small Radioisotope Power System (SRPS) is compatible with lunar missions that require a single 

Advanced Stirling Convertor (ASC) and have a low vibration requirement due to instrumentation. The 

SRPS has the potential to be adapted for other planetary missions as well. The project resulted in delivery 

of a Lunar-Derived Advanced Stirling Convertor (ASC- L) with passive balancer, simulated lunar lander 

test stand, and Single Convertor Controller (SCC). Component- and system-level testing of these items 

has been completed with plans to continue extended operation in order to gain long-term system data. The 

project produced a low mass and volume, high efficiency, and reliable SCC. The project can be adapted 

for other missions that require even less vibration than that of an ASC with passive balancer through use 

of an active balancer. Initial analysis shows that an active balancer can potentially decrease vibration 

levels beyond that of a passive balancer. 

References 

1. Dugala, G.M., et aL: Small Radioisotope Power System at NASA Glenn Research Center. Presented 

at the 9th International Energy Conversion Engineering Conference, San Diego, CA, 2011. 

2. Dugala, G.M.: Test Rack Development for Extended Operation of Advanced Stirling Convertors at 

NASA Glenn Research Center. Presented at the 7th International Energy Conversion Engineering 

Conference, Denver, CO, 2009. 

NASNTM- 2013-216561 21 


~~---- --