NASA Technical Memorandum 107252

Early Results From Solar Dynamic

Space Power System Testing

Richard K. Shaltens and Lee S. Mason

Lewis Research Center

Cleveland, Ohio

July 1996

National Aeronautics and

Space Administration

Early Results from Solar Dynamic Space Power System Testing

Richard K. Shaltens* and Lee S. Mason t

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, OH 44135

Abstract

A government/industry team designed, built and tested a

2-kW e solar dynamic space power system in a large thermal/

vacuum facility with a simulated Sun at the NASA Lewis

Research Center. The Lewis facility provides an accurate

simulation of temperatures, high vacuum and solar flux as

encountered in low-Earth orbit. The solar dynamic system

includes a Brayton power conversion unit integrated with a

solar receiver which is designed to store energy for continuous

power operation during the eclipse phase of the orbit. This

paper reviews the goals and status of the Solar Dynamic

Ground Test Demonstration project and describes the initial

testing, including both operational and performance data. System testing to date has accumulated over 365 hrs of power

operation (ranging from 400 W to 2.0-kWe), including 187

simulated orbits, 16 ambient starts and 2 hot restarts. Data are

shown for an orbital startup, transient and steady-state orbital

operation and shutdown. System testing with varying insolation levels and operating speeds is discussed. The solar

dynamic ground test demonstration is providing the experience

and confidence toward a successful flight demonstration of the

solar dynamic technologies on the Space Station Mir in 1997.

Introduction

The NASA Office of Space Access and Technology initiated

the 2-kW e Solar Dynamic (SD) Ground Test Demonstration

(GTD) Project (refs. 1 and 2). The primary goal of this project

was to conduct testing of flight prototypical components as part

of a complete SD system. Demonstrations of both system

power delivered and total system efficiency in low-Earth orbit

(LEO) were key test objectives. The SD space power system

shown in figure 1 includes the solar concentrator and solar

receiver with thermal energy storage integrated with the power

conversion unit, installed in a facility simulating an environment representative of LEO.

*Manager,Solar Dynamic Ground Test Demonstration Project.

*TestDirector, Solar Dynamic Ground Test Demonstration Project.

Programs during the past 30 years have developed SD component technologies which are now available for near-Earth orbit

applications. However, several technical challenges identified

during the Space Station Freedom Program are currently being

investigated during the GTD testing (ref. 3). These key issues are:

Flux tailoring.--integration of the concentrator and receiver

such that adequate solar flux is transferred into the cycle without

excessive flux deposition on any one area of the receiver, Control

methodology---investigate methods of varying turboalternator

compressor (TAC) speed and system thermal management in

order to maintain optimum system operation (energy management) due to large time period changes in insolation, and

Transient-mode performance.---evaluation of start-up and

shutdown transients, and multiple orbit operations.

The SD GTD project has demonstrated a complete SD

system in a thermal/vacuum environment, i.e., the large space

environmental facility, known as Tank 6, at NASA Lewis

Research Center (LeRC). The Tank 6 facility includes a solar

simulator to supply the equivalent of "one" Sun, a liquidnitrogen-cooled wall operating at 78 K (140.4 °R) which

provides a heat sink to simulate the deep space environment

(about 200 K (360 °R)), and an electric load simulator capable

of dissipating up to 4 kW of electrical power. Testing in

December 1994 has resulted in the world's first operation of a

complete SD system in a relevant environment. SD system

testing has demonstrated orbital startup, transient and steadystate orbital operation and shutdown. Flight typical components were used in the SD system wherever possible to

demonstrate the availability of SD technologies. Only the

power conditioning, control system and parasitic load radiator

were not flight hardware designs. An overview of the GTD

activities is provided by Shaltens & Boyle (refs. 4 to 6).

Solar Simulator

The LeRC solar simulator design consists of nine 30-kW

xenon arc lamps and provides a nominal flux of 1.37 kW/m 2

with a subtense angle of about 1.0 degree for testing solar
orbits.The advanced solar simulator system design results in a

50 percent improvement in the solar simulator system efficiency when compared to conventional designs. An advanced

optics system design (new collector and lens), results in the use

of only nine 30 kW arc lamps rather than nineteen arc lamps

with a conventional optics system (ref. 7). This significantly

reduces its size and initial cost as well as future operating and

maintenance costs. Fabrication, assembly, installation and

checkout of the solar simulator integrated with Tank 6 were

completed in September 1994. A detailed description of the

solar simulator design and results from early testing of a

subscale optics system are discussed by Jefferies (ref. 7), while

the initial operation is discussed by Jaworske (ref. 8).

Figure 1.--Photograph of the solar dynamic system.

Tank 6 thermal/

vacuum facility

Figure 2.--Photograph of the solar simulator and tank 6.

dynamic systems. A photograph of the advanced solar simulator (in the upper left) next to the vacuum tank is shown in

figure 2. This solar simulator provides an apparent "Sun" just

outside the vacuum tank that shines through a quartz window

into the tank to provide the desired flux (up to 1.66 kW/m 2) at

the target area. The target area (i.e., concentrator interface) is

4.79 m in diameter and 17.2 m from the apparent "Sun." A

water-cooled shutter is opened and closed to simulate various

Solar Dynamic System

The SD system includes the following major subsystems: (1)

a solar concentrator, (2) a solar receiver with thermal energy

storage, (3) a power-conversion system, (4) a waste heat

rejection system, and (5) a power conditioning and control

system. The SD system was designed to produce about 2-kW e

(at 120 Vclc) utilizing thermal energy storage with an overall

system efficiency greater than 15 percent. It should be noted

that the system performance and life were not optimized due to

the constraints of utilizing existing hardware designed for other

applications.

A block diagram of the SD system is shown in figure 3.

Energy for operating the SD system is obtained by intercepting

solar radiation by use of a parabolic solar concentrator. The

concentrator focuses the solar radiation into the cavity of the

heat receiver. The heat receiver, which incorporates thermal

energy storage, serves a dual purpose. During the on-Sun phase

of the orbit, the receiver transfers the energy to the cycle

working fluid and to the phase change energy storage material.

During the eclipse portion of the orbit, the receiver gives up

energy from the phase change material to the cycle working

fluid. The hot cycle working fluid exiting the receiver is then

s /

ELEc R 

Gs I

' VALVES kr_

I _ VALVe::, PCU

J

Figure 3.mBIock diagram of the solar dynamic

system.
expanded through a turbine thereby producing the work necessary to turn the compressor and alternator. The working fluid

passes through the recuperator (counter flow heat exchanger)

which serves to pre-heat the fluid entering the receiver, thus

increasing cycle efficiency. From the recuperator, the working

fluid is then cooled to the compressor inlet temperature by

removing the remaining waste heat energy via a gas-to-liquid

heat exchanger (gas cooler). The waste heat from the cooler is

dissipated by the heat rejection system. The fluid entering the

compressor is compressed and used to cool the alternator before

it is pre-heated at the recuperator and passed to the heat

receiver.

Design life of the GTD system is for over 1000 hr of

operation with up to 100 starts from a cold condition. The

nominal design case for the GTD was the maximum insolation

orbit with 66 min of Sun and 27 min of shade.

System Integration

Major components of the GTD system were derived from

existing designs. The TAC and recuperator came from the

Brayton Isotope Power System (ref. 9) and the off-axis concentrator, solar receiver, and radiator were based on designs scaled

from Space Station Freedom (ref. 2). Components were integrated based on the requirement that their interfaces be as

simple as possible and that their function be readily assignable

to one or another of the performing organizations. Flight

packaging was not pursued because of the desire for modularity

of components and simplification of their structural interfaces.

The modular design of the SD system also offers the potential

to evaluate advanced subsystems and components in the Tank 6

environment at a later date. Figure 4 illustrates the modular

layout of the SD components as it is configured in Tank 6.

Concentrator Subsystem

As shown in figure 1, the completed offset concentrator

structure consists of 7 hexagonal panels with 6 reflective facets

(mirrors) per panel (ref. 10). The concentrator is 4.75 m wide by

4.55 m tall and supported on a leaning tripod support structure

which attaches to a removable test stand. Facet reflectivity

exceeds 85 percent and the mass is about 2.5 kg/m 2. After

assembly, the concentrator was proof-checked on the test stand

followed by facet installation and alignment. A detailed

description of the offset concentrator design is provided by

Bahnman (ref. 11).

Receiver and Power Conversion Unit (PCU) Subsystems

Integration of the completed solar receiver with the power

conversion unit (PCU) is shown in figure 5. The receiver is used

to transfer the solar thermal energy to the cycle working fluid

and to store solar energy for system operation during eclipse.

The receiver design is essentially a scale model from the Space

Station Freedom and uses the same thermal energy storage

canister (full size) as was designed, built and tested during the

Space Station Freedom Program. Manufacturing, development

and testing of the canisters have been completed and are

discussed by Strump (ref. 12). The storage canister consists of

the Haynes 188 shell, or hollow doughnut, filled with LiF-CaF 2

eutectic phase change material. The phase change material has

a melting point of 1041 K (1873 °R). Key additions to the

canister tube design which provide enhanced performance are

heat transfer fins added to the internal flow path and a alumina

Concentrator ._ Radiators

,' : Solar

_jfp__ simulator

-- -: i:-_-'_e_=_ Liquid utilities

_ Receiver • pallet

• Power Parasitic

conditioning load

Turbo- and control unit radiator

machinery,

recuperator,

and cooler

Figure 4.mSolar dynamic system layout in tank 6.

Figure 5.mCompleted receiver integrated with the power

conversion unit.
based canister coating (ref. 13) for improved thermal emissivity. A complete description of the receiver design is provided by

Strumph (refs. 14 and 15).

The Power Conversion Unit (PCU) subsystem is a closed

Brayton cycle which includes the turboalternator/compressor

(TAC), gas coolers, recuperator, ducting and support structure.

The TAC, known as the mini BRU (Brayton rotating unit),

consists of a single stage radial flow compressor, turbine and a

brushless four pole Rice Alternator mounted on a single shaft.

Foil gas bearings are used to provide long life operation by

eliminating metal-to-metal contact of the shaft and bearings

during operation. The alternator, bearings and shaft are cooled

by the compressor discharge flow. While operating at

54 000 rpm, the TAC can produce electric power up to a

maximum of 2.2 kW (at 120 Vdc ). The PCU subsystem unit uses

a helium-xenon gas mixture with a molecular weight of 83.8 as

the working fluid. The gas mixture was optimized for heat

transfer and aerodynamic performance for the mini BRU (ref. 9).

The compressed working fluid is preheated in a recuperator by

turbine exhaust gases to increase efficiency of the cycle. A

detailed discussion of the TAC design is provided by Amundsen

(ref. 9). Acceptance testing of the PCU (known as the "hot

loop" test) was completed with demonstration of 2 kW of

electrical power. Testing was performed at local atmospheric

conditions with an electrical heater. Prior to integration of the

receiver and PCU, each assembly was covered with multilayer

foil insulation by wrapping multiple layers of nickel and

aluminum foils around the critical hot parts.

Waste Heat Rejection (WHR) Subsystem

Shown in figure 6 is the completed WHR system, which

consists of two (2) identical radiator panels plumbed in series

and a Liquid Utilities Pallet in a closed pumped liquid loop

design. The Liquid Utilities Pallet (LUP) contains the pump(s),

accumulator, sensors and an auxiliary heater for the n-heptane

coolant fluid (ref. 16). Each bonded aluminum honeycomb

radiator panel is about 1.77 m by 3.66 m with a radiating area

of 12.96 m 2. Each panel has 11 active and 11 inactive flow tubes

evenly spaced to simulate thermal transient response of a fully

redundant flow path design. Each panel is coated with a white

epoxy paint, chemglaze A276 TM, a thermal control coating. The

WHR system is integrated into the PCU loop by means of two

gas-to-liquid heat exchanger, or gas coolers. A detailed

description of the analysis, design, fabrication and testing of the

waste heat subsystem is provided by Fleming (refs. 17 and 18).

Acceptance testing of the waste heat subsystem was completed in the LeRC thermal/vacuum facilities. Both steadystate and transient operation of the WHR system was conducted.

Heat rejection during steady-state tests ranged from 2.5 to

6.3 kW t. To meet the desired fluid outlet temperature at GTD

nominal operating conditions at the lower, apparent sink temperature, radiator #2 was covered with about 1.2 m of insulation

blankets. Performance of the WHR system was as expected.

Figure 6._Complete waste heat rejection system.

Power Conditioning & Control Subsystem

The Power Conditioning and Control Unit (PCCU) contains

the power electronics. The start inverter power supply is a

commercially available, variable, controllable three-phase

power supply which provides the ability to operate the TAC

alternator as both an inductive and a synchronous electric

motor. Starting profiles are being investigated to ascertain, by

test, the optimum starting electrical characteristics. The parasitic load radiator is an integral part of the electric loop controls

and functions as an electrical sink for excess power (up to

100 percent) from the TAC which is not consumed by the user

load, accessory loads, and PCCU. The parasitic load radiator,

which is controlled by the PCCU, consists of an array of

vacuum compatible, individually controlled cal rod heaters

with enhanced emissivity characteristics (ref. 13).

The Data Acquisition and Control System (DACS) is special

test equipment whose primary function is to record system test

data. The DACS also contains the ability to communicate

setpoint conditions to the PCCU to vary speed, voltage and gain

setpoints. This allows for changing the control parameters

during the system test without the need to physically access the

PCCU within the thermal/vacuum environment.

System Operation and Testing

Integrated system testing is being conducted over the system

operating range in order to evaluate and validate previously
developed analytical models. Testing was conducted in two

phases: (1) system acceptance tests by AlliedSignal, and (2)

system characterization tests by NASA (ref. 19). Operation is

being conducted to characterize the SD system and evaluate

various analytical models over a variety of solar insolation

levels, speed conditions, orbit periods, engine inventories and

radiator variations. Further, development, verification and qualification tests are ongoing in support of the joint United States/

Russian SD flight demonstration project (ref. 20).

Flux Tailoring

Verification of the optical alignment, solar simulator to

concentrator to receiver optical interface surface was conducted in the thermal/vacuum environment of Tank 6 with the

use of a rotating flux distribution rake. The flux-distribution rake

simulated the interior cylindrical surface of the solar receiver

which allows for direct measurement of the receiver flux. Comparison of the flux test data with analytical predictions showed

excellent correlation. Further, the peak flux of 28.8 kW/m 2

compares to a worst case prediction of 42.9 kW/m 2. Special test

equipment was provided for facet alignment and flux distribution

in Tank 6 which is described by Campbell. 2J

Control Methodology

The SD system acceptance test included the concentrator,

receiver, PCU, and WHR system in the Lewis thermal/vacuum

facility with the advanced solar simulator. Acceptance testing

of the SD system has successfully demonstrated startup, transient and steady-state orbital operation and shutdown. About

2.0 kW (peak) (at 120 Vdc ) of electrical power was achieved on

February 17, 1995 while operating at 52 000 rpm (design

speed) with a turbine-inlet-temperature (TIT) of 1063.5 K

(1914.3 °R) and a compressor inlet-temperature (CIT) of

270.2 K (486.4 °R). Illustrated in figure 7 is an example of

steady-state orbital operation (over three orbits), while operating

the TAC at 52 000 rpm. The average orbital user power produced

1100-

I000-

9110-

, 2000 [

_-" 19001 RCVR EXff --_\

I \ " '_ I"

0 30 60 90 120 150 180 210 240

TIME (MIN)

Figure 7.--Data showing steady state orbital

operation,

270

was 1.83 kW e. About 107 W e of the losses are from the power

electronics contained in the PCCU. Additional accessory losses

include the cooling pump and the shutdown valves which are

estimated at 67 W e. This simulated orbit provided 66 min of

sunlight with 18 min of eclipse. Also shown is the average

canister temperature and the receiver gas exit temperature.

About 40 hr of power operation with l0 orbits including

5 successful ambient starts with 1 hot restart were accumulated

during acceptance testing. Ambient start temperature is defined

as the receiver gas temperature at 294 K (530 °R), while the hot

start temperature is the receiver gas temperature above 778 K

(1400 °R). Early evaluation of performance data showed steadystate and orbital operation of the PCU was as predicted (ref. 22).

Both thrust and journal bearings temperatures and rotor stability were shown to be within acceptable limits. During the

acceptance testing the following conclusions were reached:

(1) system starting was slower than analytical estimates because modeling ignored certain receiver mass elements which

are not critical for analysis of orbital transients; (2) the receiver

pressure drop was higher than anticipated due to incorporation

of heat transfer fins between the receiver tube and centerbody;

(3) an overall system energy imbalance existed between receiver

calorimetric calculations and solar simulator light measurements; and (4) PCCU component problems associated with

vacuum and cold environmental conditions were encountered.

Although differences were identified between the analytical

models and actual operation of individual components, the SD

system has shown to be very reliable and robust. SD system

testing performed by NASA has accumulated an additional

315 hr of power operation including 177 simulated orbits

(typically 66 rain of sunlight/27 rain of shade), and 11 ambient

(294 K (530 R)) orbital starts and 1 hot (778 K (1400 °R))

restart.

Insolation Variations

Shown in figure 8 is a test which was conducted over a

40-hr period with the TAC operating at 48 000 rpm and

illustrates an orbital startup, steady-state orbital operation and

a shutdown. Data from the integrated SD system includes: the

average receiver canister temperature, the receiver gas exit

temperature, the compressor-inlet-temperature (CIT) and the

DC power output as shown in figure 8. The solar simulator

provided four different insolation levels; 1.01, 1.06, 1.08 and

1.14 Suns (1.37 kW/m 2 -- 1 Sun), resulting in four steady-state

orbital cases, during the 93-min orbit. The initial orbit that

produces electrical power is identified as orbit 'T'. Balanced

orbital operation was achieved on orbits 4, 8, 15, and 21.

Balanced orbital operation is defined as repeatable conditions

at the same time (Sunrise or Sunset) on successive orbits,

<1.1 K (<2 °R) change in receiver gas temperatures and <5 W

change in power output. The first three cases (orbits 1-4, 5-8

and 9-15) are examples of operation of the heat receiver within
A

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Figure8.--Datashowingstartup, multiple orbitsandshutdown of the SD system.

40 45

the sensible heat regime (i.e., canister phase change material

not melted), which resulted in large temperature (137 K

(2466 °R)) and power (138 W) fluctuations. The fourth case

(orbits 16-21) is in a latent heat receiver regime (i.e., phase

change material melted), which resulted in a marked reduction

of temperature (19 K (34 °R)) and power (49 W) fluctuations

during the orbit. This is in good agreement with analytical

predictions. The TAC was operating at 48 000 rpm during the

test, except for the shutdown. Engine speed was increased to

52 000 rpm during the shutdown to speedup removal of the heat

energy in the system. The 1.01 insolation case resulted in an

overall system efficiency, Sun in to user energy, of 15 percent,

with the engine efficiency of about 26 percent. Slight increases

in both overall and engine efficiency were realized at the higher

insolation levels.

An example of data from the orbital start-up showing a

representative solar receiver heating profile is shown in figure 9.

The receiver canister temperature increases during each Sun

interval of the first three orbits until it reaches 1055 K

(1900 °R) during the third orbit. The turbine preheat requirement of 1055 K was established to overcome the potential of

compressor surge effects which were observed during the "hot

loop" testing of the PCU. It was discovered during "hot loop"

testing that if the shutdown valves were left open and the TAC

was motored, the turbine could be sufficiently preheated to

eliminate surging of the compressor. Opening the shutdown

valves which connect the compressor exit to the inlet was

1100- 2100

1000- 1900

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0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0

TIME (HRS)

Figure 9.--Data showing ambient orbital startup.

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4.5

anticipated to eliminate the compressor surge problem as these

work exactly like bleed valves used on open cycle gas turbines

for precisely the same purpose. Also shown are the receiver gas

inlet and exit temperatures which gradually increased during

the 3.5 hrs of heating.

After the canister reached 1055 K, the turbine preheat is

conducted by motoring the TAC at 30 000 rpm, with the bypass

(shutdown) valves open, for about 2 min. Note the relationship

(reversal) of receiver inlet and exit temperatures during the two

minute preheat, indicating proper flow direction. Finally, with

the bypass valves closed, the TAC is started by motoring at

36 000 rpm until self-sustained operation is observed. TAC

motoring for this start required about 4 rain.
Transient-mode Performance

Figure 10 shows a test sequence which illustrates: an orbital

startup, steady-state and transient orbital operation and a shutdown. Data from the SD system includes: average receiver

canister temperature, the receiver gas exit temperature, the

compressor-inlet-temperature (CIT), DC power output, and

TAC speed, and are all shown in figure 10. This test was

conducted over a 48-hr period with the TAC operating at

44 000, 43 000, 52 000 and 54 000 rpm. The solar simulator

provided approximately l-Sun (1.37 kW/m2), with an orbit

period of 66 min of sunlight and 27 min of shade, resulting in

27 simulated orbits producing power. Heating of the receiver

cavity required three orbits. The startup criteria is defined as the

maximum receiver canister temperature greater than 1055 K

(1900 °R). The l-Sun insolation level corresponds to about

10 kW heat to the receiver. Balanced orbital operations were

achieved on orbits 5 (@44 000 rpm), 14 (@43 000 rpm), 21

(@52 000 rpm) and 27 (@54 000 rpm). Table I summarizes the

receiver and engine performance for the four balanced orbits.

Average power output over the orbit ranged from 1.23 kW to

1.34 kW and engine efficiency (alternator output power divided

by working fluid heat input) varied from 21.5 to 26.4 percent

while overall system efficiency ranged from 13.8 to 15 percent.

Figure 11 shows the sensitivity of the receiver gas exit temperature and gas flow rate to TAC speed. Orbit 14 provided the only

example of operation in which the receiver was in the latent heat

regime.

The sequence of the speed changes was selected to effect

worst case system performance transients. The change from

43 000 to 52 000 rpm, shown in figure 10, provides insight into

the "system" response going from a hot, latent receiver to a

sensible heat receiver. This speed change is similar to what

could be expected on orbit in response to excessive receiver

energy input (e.g., extended Sun times due to higher orbit

inclinations). Figure 12 shows the short-term receiver gas

temperature and power output effects of the speed change. The

dashed lines represent the response to the speed transient and

the solid lines provide a reference of the same parameters at the

same time in the previous orbit. The data show essentially no

change in receiver gas exit temperature and a minimal (about

27.8 K (50 °R)) change in gas inlet temperature. As shown in

figure 10, the long term effect of the speed change was achieved

6 orbits later (orbit 21) upon achieving a balanced orbit at

52 000 rpm. The transition from balanced operation (orbit 14)

at 43 000 rpm to balanced conditions at 52 000 rpm resulted in

a sunset temperature (i.e., maximum orbital temperature)

decrease at the receiver gas exit of 61 K (110 °R). Similar

results were obtained for the 52 000 to 54 000 rpm speed

transient (orbit 22) with minimal short term effects and 5

transition orbits before the full impact of the speed change was

observed on the receiver temperature.

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Figure 10._Data shows TAC speed changes (44, 43, 52 and 54 K rpm) for the PCU.

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Table I.--SD system balanced orbit performance data Summary

Orbit

Parameter 5 14 21 27

Insolation, suns 0.99 1.02 1.00 0.98

ac power, kW 1.23 1.24 1.34 1.22

dc power, kW 1.18 1.19 1.28 1.17

Receiver exit, °R 1800 1842 1654 1595

Receiver inlet, °R 1542 1580 1384 1328

Compressor inlet, °R 425 420 434 435

TAC speed, rpm 44,163 43,143 52,127 54,016

Qa, s, kW 4.69 4.67 5.62 5.70

Engine efficiency, percent 26.2 26.4 23.9 21.5

Canister average, °R 1758 1802 1604 1543

Qums, kW 1.73 1.91 1.20 1.03

QRcvR, kW 9.06 9.27 9.61 9.48

Orbit efficiency, percent 14.6 14.4 15.0 13.8

10,50-

1000-

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==

900

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0.35

0.34 _

0.33 _

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40000 45000 50000 55000

TAC SPEED (RPM)

Figure 11 ...--Sensitivity of TAC speed to temperature

and flow rate.

1900 2500

1000-

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TIME (MIN)

Figure 12.--Comparison of system response to TAC

speed change (43 to 52 K rpm) vs. balanced

orbit (43 K rpm).

2000

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Initial operational and performance data has demonstrated an SD power system which is of sufficient scale and

fidelity to ensure confidence in the potential of SD technology for space. Integration of the solar concentrator and

receiver has shown that peak fluxes within the receiver

were well within worst case design predictions. System

testing has successfully shown orbital startup, transient and

steady-state orbital operation and shutdown in a relevant

space environment with a simulated Sun. Off-design thermodynamic performance data is provided which demonstrates the flexibility of the SD system under different solar

intensities and operating speeds. Over 365 hr of power

operation, ranging from 400 W to 2.0 kW e of power operation, including 187 simulated orbits, 16 ambient starts and

2 hot restart have been completed.

SD system efficiencies during orbital operation has ranged

from 13.8 to 16.1 percent. The demonstrated end-to-end

system efficiency is very good when compared to large

photovoltaic/battery systems. End-to-end orbital efficiencies of large photovoltaic/battery systems are currently

estimated to be about 4 percent for the International Space

Station. Testing to date has resulted in an improved understanding of integrated SD system operations and

performance.

Acknowledgements

SD GTD Team

The collective efforts of the SD GTD Team has resulted

in the world's first full scale demonstration of a complete

space-configured SD system in a large thermal/vacuum

facility with a simulated Sun. The authors wish to acknowledge the contributions of the SD GTD Team Members

which include: NASA LeRC, Cleveland, OH was responsible for overall project management and provided an

advanced solar simulator with the large thermal/vacuum

facility; Harris Corporation, Melbourne, FL for the offset

solar concentrator; AiliedSignal Aerospace, Torrance, CA,

for the solar heat receiver (with thermal energy storage) and

gas cooler; AlliedSignal Aerospace, Tempe, AZ, for management of the industry team members and the power

conversion system; Loral Vought Systems, Dallas, TX for

the radiator panels; and Rockwell International Corporation, Rocketdyne Division, Canoga Park, CA, for system

integration and test support. Aerospace Design & Development (ADD), Niwot, CO supplied the multilayer insulation

(MLI) for the heat receiver and power conversion subsystem while Solar Kinetics Incorporated (SKI), Dallas, TX

supplied the reflective facets for the concentrator.
References

1. Shaltens, R.K.and Boyle, R.V., "Initial Results From the Solar Dynamic

(SD) Ground Test Demonstration (GTD) Project at NASA Lewis,"

TM-107004, Proceedings of the 30th Intersociety Energy Conversion

Engineering Conference, The American Society of Mechanical Engineers, New York, NY, 1995, pp. 363-368.

2. Calogeras, J.E. and Dustin, M.O., "The Ground Testing of a 2 kW e Solar

Dynamic Space Power System," Proceedings of the 27th lntersocie_

Energy Conversion Engineering Conference, Society of Automotive

Engineers, Inc., Warrendale, PA, 1992, Vol. 1, pp. 455-460.

3. Jefferies, K.S., Ed., "Solar Dynamic Power System Development for

Space Station Freedom," NASA RP-1310, July, 1993.

4. Shaltens, R.K., "Overview of the Solar Dynamic Ground Test Demonstration Program at the NASA Lewis Research Center," NASA TM- 106876,

March 1995.

5. Shaltens, R.K. and Boyle, R.V., "Update of the 2 kW Solar Dynamic

Ground Test Demonstration Project," NASA TM-106730, Proceedings

of the 29th lntersocie_ Energy Conversion Engineering Conference,

American Institute of Aeronautics and Astronautics, Washington, D.C.,

1994, pp. 359-365.

6. Shahens, R.K. and Boyle, R.V., "Overview of the Solar Dynamic Ground

Test Demonstration Project," NASA TM-106296, Proceedings of the

28th lntersociety Energy Conversion Engineering Conference, American Chemical Society, Washington, D.C., 1993, Vol. 2, pp. 831-836.

7. Jefferies, K.S., "Solar Simulator for Solar Dynamic Space Power System

Testing," NASA TM-106393, The 1994ASME/JSME/JSESlnternational

Solar Energy Conference, (San Francisco, CA), The American Society of

Mechanical Engineers, New York, NY, March 27-30, 1994, pp. 217-222.

8. Jaworske, D.A., Jefferies, K.S., and Mason, L.S., "Alignment and Initial

Operation of an Advanced Solar Simulator," AIAA Paper 96-0102,

January 1996.

9. Amundsen, P.C. and Harper, W.B., "BIPS Turboaltemator-Compressor

Characteristics and Application to the NASA Solar Dynamic Ground

Demonstration Program," Proceedings of the 27th lntersociety Energy

Conversion Engineering Conference, 1992, Society of Automotive Engineers, Inc. Warrendale, PA, Vol. 2, pp. 239-244.

10. Schertz, P., Shabbar, S. and Lammert, L., "Development of an Improved

Facet for Space Applications," NASA CR-189109, October 1991.

11. Bahnman, D.W. and Jensen, P.A., "Design of a Solar Concentrator for the

Solar Dynamic Ground Test Demonstration Project," The 1994 ASME/

JSME/JSES International Solar Energy Conference, (San Francisco,

CA), The American Society of Mechanical Engineers, New York, NY,

March 27-30, 1994, pp. 193-203.

12. Strumph, H.J., Westelaken, B., Shah, D., Pogue, B., Chin, Jr., Alvarez, J.,

and Klucher, B., "Fabrication and Testing of the Solar Dynamic Ground

Test Demonstration Heat Receiver," Proceedings of the 29th lntersociety

Energy Conversion Engineering Conference, American Institute of Aeronautics and Astronautics, Washington, D.C., 1994, pp. 372-377.

13. de Groh, K.K., Roig, D.M., Burke, C. A., and Shah, D.R., "Performance and

Durability of High Emittance Heat Receiver Surfaces for Solar

Dynamic Power Systems," The 1994 ASME/JSME/JSES International

Solar Energy Conference,(San Francisco, CA), The American Society of

Mechanical Engineers, New York, NY, March 27-30, 1994,

pp. 251-264.

14. Strumph, H.J., Avanessian, V., Ghafourian, R., and Huang, F.J., "Thermal

and Structural Analysis of the Heat Receiver for the Solar Dynamic

Ground Test Demonstrator," The 1994 ASME/JSME/JSES International

Solar Energy Conference, (San Francisco, CA), The American Society of

Mechanical Engineers, New York, NY, March 27-30, 1994,

pp. 223-234.

15. Strumph, H.J., Krystkowiak, C. and Killackey, J.J., "Design of the Heat

Receiver for the Solar Dynamic Ground Test Demonstrator Space Power

System," Proceedings of the 28th Intersociety Energy Conversion Engineering Conference, American Chemical Society, Washington, D.C.,

1993, Vol. 1, pp. 469-475.

16. Fleming, M.L., and Flores, R.R., "Radiator Selection for Space

Station Solar Dynamic Power Systems," Proceedings of the 22nd

IntersocieD' Energy Conversion Engineering Conference, American Institute of Aeronautics and Astronautics, New York, NY, 1987,

pp. 208-213.

17. Fleming, M.L., Flores, R.R. and Sharpe, R.R., "Solar Dynamic

Ground Test Demonstration Radiator Design and Test," Proceedings

of the 29th Inter-society Energy Conversion Engineering Conference,

American Institute of Aeronautics and Astronautics, Washington, D.C.,

1994, pp. 378-383.

18. Fleming, M.L. and Flores, R.R., "Solar Dynamic Radiator Design

Development," The 1994 ASME/JSME/JSES lnternational Solar Energy

Conference, (San Francisco, CA), The American Society of Mechanical Engineers, New York, NY, March 27-30, 1994, pp. 245-250.

19. Mason, L.S. and Kudija, C.J., 1994, "Solar Dynamic Ground Test Demonstration System Test Plans," The 1994 ASME/JSME/JSES International

Solar Energy Conference, (San Francisco, CA), The American Society of

Mechanical Engineers, New York, NY, March 27-30, 1994,

pp. 175-184.

20. Huckins, E. and Ahlf, P., "Space Station Power Requirements and Issues,"

Proceedings of the 29th Intersociety Energy Conversion Engineering

Conference, American Institute of Aeronautics and Astronautics, Washington, D.C., 1994, pp. 608--612.

21. Campbell, J.S. and Jensen, P.A., "Design, Analysis and Test of a Solar

Concentrator for Space Applications," The 1994 ASME/JSME,/JSES

International Solar Energy Conference, (San Francisco, CA), The American Society of Mechanical Engineers, New York, NY, March 27-30,

1994, pp. 205-216.

22. Mock, T.A., "Solar Dynamic Ground Test Demonstrator (SDGTD) System Orbital and Startup Control Methods," The 1994 ASME/JSME/JSES

International Solar Energy Conference, (San Francisco, CA), The American Society of Mechanical Engineers, New York, NY, March 27-30,

1994, pp. 185-191.
REPORT DOCUMENTATION PAGE I Form OMBNo. Approved 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

July 1996 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Early Results From Solar Dynamic Space Power System Testing

6. AUTHOR(S)

Richard K. Shaltens and Lee S. Mason

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

WU-233-03-0B

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-10176-1

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM- 107252

11. SUPPLEMENTARY NOTES

This paper is also published in The Journal of Propulsion and Power, American Institute of Aeronautics and Astronautics, Washington D.C., Vol 12, No. 5, Sept.--Oct. 1996. Responsible person, Richard K. Shaltens, organization code

5490, (216) 433-6138.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 20

This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A government/industry team designed, built and tested a 2-kW e solar dynamic space power system in a large thermal/

vacuum facility with a simulated Sun at the NASA Lewis Research Center. The Lewis facility provides an accurate

simulation of temperatures, high vacuum and solar flux as encountered in low-Earth orbit. The solar dynamic system

includes a Brayton power conversion unit integrated with a solar receiver which is designed to store energy for continuous

power operation during the eclipse phase of the orbit. This paper reviews the goals and status of the Solar Dynamic

Ground Test Demonstration project and describes the initial testing, including both operational and performance data.

System testing to date has accumulated over 365 hrs of power operation (ranging from 400 watts to 2.0-We), including

187 simulated orbits, 16 ambient starts and 2 hot restarts. Data are shown for an orbital startup, transient and steady-state

orbital operation and shutdown. System testing with varying insolation levels and operating speeds is discussed. The solar

dynamic ground test demonstration is providing the experience and confidence toward a successful flight demonstration of

the solar dynamic technologies on the Space Station Mir in 1997.

14. SUBJECT TERMS

Space power; Solar dynamic; Brayton cycle; System testing

17. SECURITY CLASSIFICATION

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18. SECURITY CLASSIFICATION

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Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

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11

16. PRICE CODE

A03

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