Dr. Kitt Reinhardt

1
Enabling Solar Cell and Solar Array Technologies
for Next-Generation DoD, Commercial, and NASA
Missions
Dr. Kitt Reinhardt Space Vehicles Directorate Air
Force Research Laboratory Kirtland AFB,
NM kreinhar_at_pop500.gsfc.nasa.gov
2
AFRL and Space Power Groups
CUSTOMERS
AF SPACE MISSILE SYSTEMS CENTER
SBL
AIR FORCE RESEARCH LABORATORY (GEN NIELSEN)
Work- force 5700
Budget AF 1.3B, Customer 1.1B
MATERIALS MANUFACTURING DIRECTORATE (ML)
DIRECTED ENERGY WEAPONS DIRECTORATE (DE)
SPACE VEHICLES DIRECTORATE (VS) (COL MCCASLAND)
MUNITIONS DIRECTORATE (MN)
INFORMATION DIRECTORATE (IF)
PROPULSION DIRECTORATE (PR)
SENSORS DIRECTORATE (SN)
Classified
AIR VEHICLES DIRECTORATE (VA)
SPACECRAFT TECHNOLOGY DIVISION (VSS) (KIRT MOSER)
INTEGRATED EXPERIMENTS DIVISION (VSE)
SPACECRAFT COMP. TECH BRANCH (VSSV) (STEVE
HUYBRECHTS)
SPACE POWER GENERATION (Reinhardt/Senft/ Mayberry
)
spacecraft technologies
SPACECRAFT MECHANISMS
CONTROL SYSTEMS
LARGE DEPLOY OPTICS
INTEGRATED STRUCTURAL SYSTEMS
AEROSPACE POWER DIVISION

• PMAD
• storage
• thermal

PMAD Electronics
Ther- mal
Designated an AF Advanced Technology Demonstration
Program (ATD)1-of30 in AFRL
Solar Cell/Arrays

Energy Storage
Pay- Load
Funded devt of 18.5, 23 2-J, 24-27 3-J, 35
4-J solar cells
3
Outline

• Solar Cell Fundamentals
• - Device Physics and Efficiency
• - Available Space Solar Cells
• Space Solar Arrays
• - Basic Configurations
• - Solar Array Design Drivers
• - Array Sizing and Mass Calculations
• - DOD, Commercial and NASA Requirements
• Near-Term Solar Cell/Array Design Options
• ? 28-30 3-J cells, advanced light-weight
  arrays (gt150W/kg)
• Next-Generation Solar Cell/Array Design Options
• ? 35 4-J cells, thin-film solar cells/arrays
  (gt400W/kg)

4
Solar Cell Fundamentals

• Device Physics and Efficiency
• Available Space Solar Cells

5
Solar Cell Basics
The solar photovoltaic cells, aka solar cell,
is based on the photovoltaic effect
appearance of a voltage across an illuminated
semiconductor p-n junction
In space
1350 W/m2
solar array
Si (10) ? 135 W/m2 3-J(30) ? 405 W/m2
6
Terrestrially
Solar energy striking earths surface in 1 minute
exceeds the total energy consumed by worlds
population in 1 year

• 10m2 _at_ 5 500W
• today PV modules 3/W
• full PV system (PV PMAD)
• 0.25 kWh (1/3 PV, 1/3 PMAD, 1/3 other)
• 3-4x conventional utilities

• Annual total U.S. electricity
• demand is about 3 x 1012 kWh
• Incident solar energy
• 6.40 kWh/m2/day (annual
• daily average) for the SW US.
• U.S. demand 10,000 mi2 in SW
• or 271 mi2 per state for
• - 10 cells
• - fixed-plate array
• - land area is 2x array area

7
Space Solar Cells
3-Junction Solar Cell
1-Junction Solar Cell
Top Grid Fingers
5-7 mil
p-type emitter
Bottom Planar Contact
n-type base
25-30 efficient (25-30cm2)
12-19 efficient (20-70cm2)
GaInP/GaAs/Ge
8
Single-Junction Solar Cell Operation
bulk Si
Si
h?
h? gt Eg(bandgap) ? electron (e) hole (h) pairs
(bond-energy) via h?
absorption h? lt Eg(bandgap) ? h? transmission
e
h
Si
Si
Si
Si
solar spectrum 1350 W/m2
1.1?m
Eg(eV) cut-off ?(?m) Si
1.1 1.1 GaAs 1.4
0.87 GaInP2 1.9 0.65 Ge
0.67 1.85
?(?m)
0.2 1.0 2.0
p-type Si
absorbed
n-type Si
cut-off ?(?m) 1.24/Eg(eV)
solar cell
transmitted
9
Single-Junction Solar Cell Operation
Gp(EHP/cm3-sec)
h
p
e
Ln

VL
p/n junction
IL
-
h
Lp
n
e
e
Gn(EHP/cm3-sec)

• IL qA(GpLn GnLp), decreases w/increasing
  Eg
• VL lnIL 2/3 Eg
• Power VL x IL

10
Calculation of Efficiency
Solar Cell Light I-V Curve
11
Single-Junction Solar Cell Efficiency vs Bandgap
Its noted that solar cell efficiency is higher
under the terrestrial spectrum than space
spectrum more wasted UV and IR in space
12
Multijunction Solar Cells
Single Junction Solar Cell Losses
3-Junction GalnP2/GaAs/Ge Cell Losses
J1 J2 J3
AM0 Solar Spectrum (1350 W/m2)
0.87 um
1.85 um
Spectral Irradiance (Wcm-2 mm-1)
Spectral Irradiance (Wcm-2 mm-1)
0.67 um
1.1 um
e -
Wavelength (um)
Wavelength (um)
e -
e -
e -
e -
LOSSES 29 - h? gt Eg
19 - hv lt E g 19 - Eg gt qVoc
7 - reflection/grid lines 5 -
quantum efficiency 4 - fill
factor Si (Eg 1.1 eV) 16 efficient
e -
Energy Lost to Lattice As heat
e -
e -
Energy Lost to Lattice As heat
Eg
J1
J2
J3
Eg 1.85 eV GaInP2
Eg 1.42 eV (GaAs)a
Eg 0.67 eV (Ge)
Available energy 35 17
31 of total spectrum Cell efficiency
15 10 2 27 efficient
Additional p/n junctions minimize
super-bandgap (heat) and sub-bandgap
(transmission) photon losses
13
Multijunction Solar Cells

• Requirement semiconductor lattice and current
  matching
• limited number of matched material systems
• theoretical 3-J design 50 efficiency w/Eg
  1.75, 1.18, 0.75 eV
• practical 3-J design w/Eg1.88, 1.43, 0.67 eV
  30-31 efficiency
• practical 4-J design w/Eg1.88, 1.43, 1.05, 0.67
  eV 36-38 efficiency
• theoretical 30-J design 70 efficiency (same
  as thermodynamic limit)

AF/SNL patent
14
Solar Cell AM1.5 and AM0 Efficiencies - highest
reported -
Date Progress in Photovolt Res App, 2001
9287-393, Keith Emery (NREL), Shiela Bailey
(NASA GRC)
15
Commercially Available Space Solar
Cells (Lot-average efficiency, bare cell, AM0 -
1353W/m2 _at_ 28?C)
14 Si, 19 1-J, 2-J are obsolete, 17 Si and
gt26 are SOTA
16
Space Solar Arrays

• Basic Configurations
• Solar Array Design Drivers
• Array Sizing and Mass Calculations
• DOD, Commercial and NASA
• Unique Requirements

17
Solar Array Configurations
Increased launch capability ? increased
spacecraft capability ? increased power
demand ? increased array size ? arrays have
evolved by necessity

• single mult-panel
• flip-out (paddles)
• large cylindrical

• kinematically complex
• multi-panel deployment
• innovative structural platforms
• unique deployment systems
• novel mechanisms
• higher efficiency photovoltaics

simple body-mounted solar panels
10s 100s Watts - 1958 Vanguard I - 2001
University experimental sats
100s 2000 W today
High-Power 1-25kW
18
Satellite Power Trends
19
Space Solar Array Configurations - Deployed Array
-

• 1-4 wings per spacecraft, 1- or 2-axis
  stabilized
• low-high power level 100sW 25kW (110kW for
  ISSA solar array)
• array structural platforms
• - rigid Al honeycomb or tensioned flexible
  blankets
• deployment hinges or articulated mast

3
1
2
4
CIC
1 solar cell assembly solar cell,
interconnect, coverglass (CIC) 2 panel assembly
solar cell assembly, wiring harness, panel
substrate, hinges 3 array structure (3) yoke
structure assembly, (4)root hinge mechanism
assembly, (5)deployment synchronization system
(cable/pulley/pantograph/etc), (6)tie down
system w/release
5 6
stowed
20
Space Solar Array Configurations - Body-Mounted -

• experimental / low power comsats
• - 10s - 2000 W
• simple no hinges or yoke
• generally spin-stabilized

Boeing 376 telescope cylindrical design
800-2000W
1 solar cell assembly solar cell,
interconnect, coverglass (CIC) 2 panel assembly
solar cell assembly, wiring harness, panel
substrate 3 array structure spacecraft
deployment synchronization system
(springs/cable/pulley/etc), tie-down system
w/release
stowed
21
Solar Array Design Drivers
22
Impact of Increased Cell Efficiency
Solar Cell
Efficiency Decrease array area, mass stowed
volume and cost/Watt, or
increase power _at_ constant array size.

• reduced stowed volume
• less sensor obscuration
• smaller COP/COM
• less attitude control fuel

Array Area Reduce array mass, stowed volume,
drag, ACS fuel budget,
Center of Pressure/Mass
Array Mass Increase payload mass, reduces
launch costs.

• reduced launch costs
• increased payload mass
• reduced attitude and
• station keeping control

Stowed Volume Completely enabling for future
EELV having significantly reduced fairing volume
compared to Titan IV launch vehicle. ? Titan IV
w/ 99 m3 fairing replaced by EELV w/ 64 m3
fairing
or

• enable more payload power
• enable solar cell retrofit using existing panel
  design
• address power vs stowed volume challenge

Array Power for constant array area
23
Impact of High Efficiency Cells
DMSP 500 nmi, 90 inclination

• - asumptions 5 kW EOL (1E14 e/cm2 1 MeV, 6 mil
  coverglass, 60C)
• - LEO10K/Kg, GEO65K/Kg
• Solar Array Performance
• 38 W/kg (EOL), 43 m2, 132Kg
  0 -
  - 5000 W
• (state-of-practice 14 Si)
• 46 W/kg (EOL), 34 m2, 109Kg
  23 230K / 1.5M
  9 6407 W
• (state-of-practice 17 Si)
• 72 W/Kg (EOL), 19 m2, 69Kg
  63 630K / 4.1M
  24 11223 W
• (state-of-art 27 3-J)
• 90 W/Kg (EOL), 15 m2, 56Kg
  76 760K / 4.9M
  28 14660 W
• (next-generation 35 4-J)

array mass launch cost
area savings (Kg) savings
() delta
(LEO / GEO) (m2)
same mass area w / more power
or

• redcuced launch costs
• increased payload mass
• reduced attitude and
• station keeping control

• reduced stowed volume
• less sensor obscuration
• smaller COP/COM
• less attitude control fuel

• solar panel retrofit
• more payload power
• reduce power vs
• shroud problem

mission enablers
24
Impact of High Efficiency Multijunction Solar
Cells

• GPS IIF
• Sats 1-6 (originally)
• 3 panels/wing (6 panels)
• total of 5 panels w/solar cells
• 12.3 BSR Si solar cells
• EOL total power 1646 W
• GPS IIF Modernization
• up to 12 sats and adding L-Band
• 2 panels/wing (4 panels)
• total of 3.4 panels w/solar cells
• 24.5 3-J solar cells
• EOL total power 2405 W

Trade Study for a 4 kW DOD Satellite (circa 1998)
30 reduction in array mass along with 45 more
power
Increased solar cell efficiency enables greater
spacecraft payload mass and power budgets and/or
reduced spacecraft mass and launch costs
25
Impact of Next-Gen PV and Battery on Satellite
Payload
Intelsat 8AGEO Communications
Advanced Power
State-of-Practice
963
1341
Balance of S/C
35 Multjunction Solar Arrays Lithium Polymer
Batteries
Balance of S/C
Fuel
1788
1788
Fuel
Power
Pwr
632
Atlas IIAS 39 Increase in Bal of S/C 33
increase in power 7.3 kW Bus (EOL) 9.5 kW Array
(BOL)
Prop
Prop
Atlas IIAS Launch (3570 kg Initial Mass) 5.5 kW
Bus (EOL) 7.2 kW Array (BOL)
254
18.5 PV NiH2
187
187
State-of-Practice
Advanced Power
GlobalstarLEO Communications
94
94
Fuel
Fuel
13
Balance of S/C
Prop
Balance of S/C
35 MJ Solar Arrays Flywheel Energy Storage
224
13
Prop
Pwr
120
Power
27
224
18.5 PV NiH2
357 kg Initial Mass (21 Reduction) 5 S/C per
Launch
450 kg Initial Mass 4 S/C per Launch
26
General Solar Array Sizing
must understand solar array sizing to do trades

• Approach
• 1) Pick orbit (altitude, inclination,yrs)
• 2) calculate
• Radiation fluence (1 MeV e- equiv)
• Operating temperature
• BOL efficiency (28C)
• EOL efficiency (operating temp)
• blanket area and mass
• array area (length,width) and mass
• total mass
• CoP
• CoM

• solar cell type
• coverglass thickness
• array design
• panel
• support structure

vs
27
Solar Array Sizing Basics
Ref Wertz and Larson

• Solar Array Design Process
• Determine mission requirements and constraints
• - mission lifetime
• - average power required at EOL during daylight
  and eclipse
• - spacecraft orbit parameters sun incident
  angles, eclipse periods,
• shadowing, radiation environment
• Calculate power requirement
• - determine solar array power (PSA)
  required during daylight to
• power spacecraft during entire orbit

Pd - power reqt during daylight Pe - power
reqt during eclipse Td - daylight time (64min
for LEO) Te - eclipse time (35min for LEO) Xd
- SA to load eff. due to regulation and
batt charging losses (0.8-0.85) Xe - eff.
due to regulation and batt discharging
losses (0.6-0.65)
28
Solar Array Sizing

• Estimate ideal solar array power output (Po) in
  W/m2 for
• for the sun (1353 W/m2) normal to the
  array

Ideally Si cells at 17.0 efficiency ? Po
230 W/m2 3-J cells at 27.0
? Po 365 W/m2

• Estimate practical solar array power output (Pp)
  in W/m2
• accounting for inherent losses and sun
  incident angle

In practice must account for solar cell and
panel losses
29
Solar Array Sizing
PP PoIdcos? cos? is the cosine loss, ? is the
Sun incident angle between the normal to the
array and the Sun line At the beginning-of-life
(BOL) PBOL PoIdcos?
flat (deployed) solar array, ? will vary
throughout mission

• Estimate solar array on-orbit life degradation
  (Ld) due to
• space radiation trapped energetic electrons
  and protons in
• Earths Van Allen belts degrade solar cell
  voltage and current
• - e.g. Si arrays in low LEO orbits degrade
  2.5/yr, GaAs and MJ 1.5/yr
• thermal cycling in/out eclipse, micrometeoroid
  strikes, thruster
• plume impingement material outgassing, UV
  degradation
• - e.g. Si and GaAs arrays in LEO orbits degrade
  1.25/yr

Ld (1- degradation/yr)satellite life
30
Solar Array Sizing

• Solar array end-of-life power (PEOL) and area
  (ASA)
• PEOL PBOLLd
• ASA PSA / PEOL

• Notes
• for 2-axis sun-tracking the array area A ASA
• body-mounted array on cylindrical spinner array
  A ?ASA
• trade required between ASA and complexity of
  stabilizing
• spacecraft with dive systems, also many array
  shapes

31
Temperature Effects

• solar cell temperatures LEO -120 - 80C,
  GEO -180 - 80C

• solar cell have negative temperature
  coefficients
• typical values Si(-0.055/C),
  GaAs(-0.025/C), 3-J(-0.045 /C)
• Concentrators, Sun/Mercury and LILT
  (low-intensity-low-temp) missions

32
Radiation Effects
Energetic electrons and protons trapped in
Earths Van Allen Belts collide with solar cell
semiconductor lattice producing defects which
enhance electron and hole recombination ? reduce
Vm and Im

• General approach to determine degradation
• Define solar array orbit (altitude,
  inclination) and lifetime
• Determine equivalent 1 MeV electron fluence due
  to electrons and protons
• - JPL Solar Cell Radiation Handbooks for Si
  and GaAs
• - Aerospace Corp/Air Force Report for
  3-Junction Cells
• - Space Radiation Software Programs
• Use solar cell degradation data for 1 MeV
  electrons measured in laboratory
• - JPL Solar Cell Radiation Handbook and
  Aerospace/Air Force Report
• ? Solar cell current, voltage degradation
  coefficients

33
Data from JPL GaAs Solar Cell Radiation Handbook

• proton tables
• trapped
• solar flare

34
Radiation Degradation

• Vendors provide values for solar cell Imp/Impo,
  Vmp/Vmpo and Pmp/Pmpo
• versus 1 MeV fluences (e/cm2) (typically 1E13
  - 1E15)
• EOL efficiency versus fluence also common

Emcore Data
35
10 Year Mission Comparisons
3-J Cell Aerospace Corp./Air Force Report (Dean
Marvin)
In general, radiation resistance 3-J gt 2-J gt
GaAs gt Si
36
Solar Array Area vs Solar Cell Type for 5 kW
(EOL) Flat-Deployed in LEO - DMSP 500 nmi, 90
inclination -
Assumptions 1) Array Area (ASA) PSA / PEOL,
where PEOL (1350)x(cell EOL Eff _at_
60C)x(Loss factor) 2) EOL 1E14 e/cm2 _at_1MeV
10 yr LEO (500nmi, 90 incl.) w/12 mil coverglass
(DMSP) 3 yr LEO (500nmi, 90 incl.) w/6 mil
coverglass 3) Loss factor 0.80 (panel
assembly, thermal cycling, out-gassing,
micrometeoroid, UV degradation) 4) Pin 1353
W/m2, no shadowing, cos? 1
37
Array Mass Calculations
Flexible Fold-Out Array
Rigid Fold-Out Array
cell assembly mass
cell assembly mass
sub (flex blanket) hinges
wiring circuit / harness mass
panel hinges wiring
harness mass
blanket assembly
panel assembly


blanket housing assembly mast, mast
canister, actuator motor blanket deployment
assembly pallet/lid structure, blanket
preload/ release mechanism
structure mass yoke, root hinge, deployment
sync. system, tie down system w/release
coilable mast
solar array mass
solar array mass
38
Solar Cell Laydown

• solar cells come in all sizes and shapes

• series and parallel cell strings
• build-up voltage and current
• bypass diodes for shadow mitigation

I ? Ii
17

24-28
Ii
I1
V1

13 ISSA
V ?vi

12-14

24-28
_
Vi

• varying cell voltages and currents
• Si Jmax 30-40 mA/cm2
• Vmax 0.5-0.6 V
• 3-J Jmax 16 mA/cm2
• Vmax 2.3 V

Vmax 30-160V arrays in space P VI 10s
1000s W (expt missions) 5-25kW
(comsats) 110kW ISSA
39
Panel Assembly Component Mass -
Rigid Fold-Out -
Coverglass/Interconnect
Bare Cell Mass
Table 1
Table 2
40
Panel Assembly and Support Structure Mass - Rigid
Fold-Out -
Substrate/Hinges/Wiring (SHW)
Table 3
Panel Assembly Mass (PAM) bare cell
coverglass/ interconnect substrate/
hinges/wiring
Array Structure Mass (ASM)

• yoke
• root-hinge
• deployment
• tie-down/release
• contingency

assumption ASM 0.3 x PAM
good approximation

41
5 kW (EOL) Solar Array Mass vs Cell Type
Mission 500 nmi, 90incl, sun-tracking, 3 yr, 6
mil coverglass (1E14 e/cm2 1MeV)
Data from Tables 1-3 assuming solar cells
occupy 90 total array area.
42
SOTA Rigid Arrays
BSS601 Array (8-10kW) 55-60 W/Kg BOL w/27 3-J
cells
GPS IIF PUMA Array (Able-Eng) BOL 3 kW, 45
W/Kg w/24.5 cells
NASA SWIFT PUMA Array (Able-Eng) BOL 2.7 kW,
56 W/Kg w/25/21 cells EOL 2.1 kW, 45 W/kg
43
Flexible Fold-Out Arrays
Advanced Photovoltaic Solar Array (APSA)
(TRW/JPL Final Report, Nov 94)
State-of-Practice Flight Design ( TERRA
heritage )
44
5 kW (EOL) Solar Array Mass vs Cell Type - Flex
Fold-Out -
Mission 500 nmi, 90incl, sun-tracking, 3 yr, 6
mil coverglass (1E14 e/cm2 1MeV)
Milstar lt 30 W/Kg BOL
w/18.5 cells
TERRA 44W/Kg
w/28 cells
in practice reqts on stiffness,
reliability, and under-optimization result in
heavier design - but - a low power
flexible array design has been qualified at gt
100 W/Kg
Data from Tables 1, 2,4 assuming solar cells
occupy 90 total array area. Assume infinite
backshielding arrays would be 5-10
larger/heavier
45
ISSA Flexible Solar Array
(b)
30 W/Kg BOL

• Final Proposed Configuration
• 110 kW per Goldin
• 262,400 Si cells (8cm x 8 cm 14.2)
• Likely de-scaled to 80KW
• 40kW to scientist after battery
• charging, PMAD, thermal losses

46
Rigid/Flexible Solar Array Comparison
Rigid Array Advantages

• Have heritage and reliability in favor
• Smaller than flexible arrays due to backside
  shielding
• Lighter than flexible arrays flown to date
• Can generate power in stowed configuration -
  flex-array cant
• Flexible arrays enabling for some missions
• 2x reduced stowed volume to fit between
  spacecraft and fairing
• Wider/shorter array configuration reduces COP
• ? reduces reaction wheel and torque rod size,
  mass, cost
• Flex array has greater proportion of its mass
  at spacecraft
• interface (canister and blanket box mass) ?
  COM closer to sc
• TERRA selected a flexible array design due to,
  in part,
• stowage, COP, COM considerations

Flexible Array Advantages
47
DoD, Commercial, and NASA Unique Solar Array
Requirements
48
Solar Array Space Environment
Military, Com, NASA Code Y
NASA Code S

• high solar
• intensity
• high temp

• low-intensity, low-temp
• (LILT)
• high radiation-fields
• dust on Mars

• electrons,
• protons,
• atomic oxygen
• charging effects
• (field/particle
• measure missions)

49
Unique Solar Array Requirements

• Military (DOD)
• - LEO,MEO,GEO reconnaissance / intelligence
  / communications / weapons
• - SmallSats (lt1kW) ultralow mass and cost
  ? thin-film PV and arrays
• - LargeSats (1-30kW) GPS, SBIRS, Adv EHF,
  WGS, classified maximize
• payload mass and power and array W/m2 and
  W/m3 ? 3 4-J cells
• - MonsterSats (30 100skW) high power
  platforms,
• space-based electric laser needed 100s
  kW MW ? thin-film PV and arrays
• - Reliable /survivability to natural and
  man-made threats
• MEO orbits radiation/lifetime issues
• laser high-temp cells, small arrays
• pellet small, possibly maneuverable
  
• detection small radar cross-section
• - Cost not greatest driver ? deploy SOTA
  and next-generation solar arrays and
• payloads first flight of all III-V 1J,
  2J, 3J solar cells on military spacecraft

50
Unique Solar Array Requirements

• Commercial (Comsats) - LEO, MEO, GEO
• - Greatest driver is generating revenue
• - focused on solar array that yields lowest
  /W at mission level
• due to ultra-competitive nature of
  communications business
• a) minimize array cost ? very interested in
  thin-film PV arrays
• b) maximize array W/Kg ? maximize payload
  power and mass
• ? i.e., payload revenue 740K/Kg-yr
  for transponders
• c) power scalability (20-30kW) for more
  transponders
• ? higher efficiency solar
  cells rule the day
• ? ? 3-J cells used exclusively
  by Boeing (Hughes)
• and nearly exclusively by
  Loral and LM

51
Unique Solar Array Requirements

• Science (NASA) earth science,
  interplanetary, interplanetary surface
• - Earth Science low-mid power (lt7-8kW)
  focused on ultra-reliability, generally
• conventional array designs have been
  acceptable. Some missions require SOTA
• ? SWIFT using 2- and 3-J cells
• - Interplanetary / Interplanetary Surface
• (see NASA OSS Report on Solar Array
  Technology POC Ed Gaddy)
• - reduced solar array mass and stowed
  volume is critical
• - desire lightweight solar arrays
  that stow in compact launch volumes
• ? advanced more costly arrays
  designs demod gt 100 W/Kg flex array design
• - high power/high-voltage arrays for
  solar electric propulsion
• - Special Environments
  
  Approaches
• High-intensity / high-temperature (Mercury and
  Solar missions)
• Low-intensity / low-temperature (LILT) (beyond
  Mars missions)
• High radiation fields (Europa, Jupiter)
• Electrostatically clean arrays for fine
  magnetic measurements
• Dust on Mars

See OSS report, But basically better solar
cells and coatings
52
Near-Term Solar Cell/Array Designs

• Multijunction Solar Cells
• Lightweight Array Designs

53
AFRL Funded Space Solar Cell Devt
- AFRL leading government multijunction cell
development -
54
Spectrolab Solar Cells
Solar Cell Product Insertion Roadmap

• 26.5 Improved Triple Junction (ITJ) in
  Production Since November 2000
• Over 120,000 ITJ cells built (Average 26.8 _at_
  2.230V, 27.0 _at_ Max Power)

TJ GaInP/GaAs/(Ge)
Improved
Ultra TJ
TJ
DJ GaInP/GaAs/(Ge)
SJ GaAs/(Ge)
GaInP
GaInP
GaAs
Ge
GaAs
GaAs
Ge
Ge
Ge
2004
1995
1999
2001
2002
1997

Min Avg BOL
28.0
26.5
33.0
21.5
19.0
24.5
0.90V
2.05V
_at_ Load Point
2.23V
2.27V
2.69V
2.22V
Avg EOL (5E14)
19.1
21.8
23.3
25.5
15.4
30.4
AM0 Efficiency (28oC)
55
Spectrolab Solar Cells
UTJ IRD Program Goals

• Improved BOL Performance (AM0 _at_ 28oC)
• - 28.0 Minimum Average Efficiency _at_ VLOAD
• Preliminary Cell Parameters
• - Vmp 2.320V
• - Jmp 16.5 mA/cm2
• - Eff _at_ Pmp 28.3
• - VLOAD 2.270V
• - Min Average I _at_ VLOAD 16.7 mA/cm2
• Improved EOL Performance (1 MeV electrons)
• - NPmp 0.91 _at_ 5E14 (25.5) and 0.87 _at_ 1E15
  (24.4) - NVmp 0.97 _at_ 1E14, 0.94 _at_ 5E14
  and 0.92 _at_ 1E15
• Must incorporate bypass diode protection that
  enables
• lt140 micron cell thickness (flat back, no RTC)

56
Spectrolab Solar Cells
Preliminary UTJ Performance Data
Ave. Voc 2.695 V Ave. Jsc 17.40 mA/cm2 Ave.
Eff 29.0_at_ Pmp Ave. FF 83.6 Ave. Vmp 2.380
V Cell area 4.00 cm2
57
Spectrolab Cell Costs
Satellite Benefits - Price/Power
From SJ to 4J, Spectrolabs Plan cuts GaAs cell
prices in half.

398
326
270
263
206
236


SJ
DJ
ITJ
UTJ
4J
TJ
Assumes 5E14 e-/cm2 AM0 135.3 mW/cm2 28C
58
Spectrolab Panel Costs
59
EMCORE 3-J Solar Cells

• EMCORE 3-Junctions Solar Cells
• cell cost 250-300/W in large quantity
• panel cost targeted in at lt500/W in 02

2,500 production space cells (27.5cm2) min. ave
efficiency 27.5 in Jan 02.
Navid S. Fatemi November 29, 2001

60
Solar Cell/Array Cost Trades

• 3-J 27 solar arrays slightly more expensive
  than Si arrays
• - panel level Si 125-150/W, 3-J 450-500/W
  in large quantity
• - cost has come way down but includes no extra
  engineering or qualification
• Results of Trade Study by Gene Ralph of Tecstar
  Inc, entitled Solar Cell Array System
• Trades, 37th AIAA Aerospace Sciences Meeting,
  Jan 1999

Note costs do not include engineering or
special qualification
Assumptions GEO 20 East, 15yrs, coverglass
thickness 4mil, front-and backside shielding,
Equinox, AM0, packing factor 0.90. LEO 1400km,
45 Inc., rest same
The above does not consider cost savings of
lighter/smaller array! Assuming 11K/kg for LEO
launch, an array area cost penalty (attitude
control fuel) of 8K/m2 for LEO mission, the 3-J
array is 25 and 35 cheaper than the 17 Si and
12.6 Si array, respectively.
61
A Glimpse Into Future Array Designs
- Courtesy of Able Engineering -
UltraFlex
Scarlet
CellSaver
conventional PUMA
ORP
SquareRigger
www.aec-able.com Tel [phone] Fax
805.685.1369
Able Engineering Company Corporate Headquarters,
7200 Hollister Ave., Goleta, CA 93117
    
62
The UltraFlex Solar Array

• Mission-Enabling Technology for Mass and Stowage
  Critical Applications
• Ultra-Lightweight Extremely Low Stowage Volume
• 1/4 to 1/3 the weight of standard arrays
• over 150 W/kg BOL with 27 TJ solar cells, 200
  W/kg BOL with 35 4J cells,
• and 300 W/kg BOL with 10 thin film PV
• lt1/4 the stowed volume of standard arrays
• Allows spacecraft to maximize payload or reduce
  launch costs
• Routinely evaluated as the most mass efficient,
  lowest risk, array solution for interplanetary
  missions

Typical Application Mars 2001 Lander (LMA)
UltraFlex Qualification Wing
ABLE Patent 5,296,044
Advanced Applications CNSR, INSIDE Jupiter,
Solar Probe, Pluto Kuiper Belt
63
CellSaver Overview

• CellSaver is a simple on-panel concentrating
  element
• Collapsible, self deploying, 2X reflective
  elements
• Ultra-thin, one-piece titanium construction, with
  space-compatible reflective coating
• Replaces every other row of cells with simple,
  low mass, low cost reflectors
• Adaptable to rigid panel or flexible blanket
  systems
• Reduces cost mass using heritage systems
• Standard large area cell, interconnect and
  laydown
• Standard panels and stack height spacing
• Standard deployment synchronization mechanisms
• Compatible with single-axis tracking
• Wide off-pointing acceptance a 10 (before
  gradual roll-off), b 30 (cosine roll-off)
• Operationally friendly off-pointing
  characteristics
• Minimal re-qualification for rigid panel retrofit
  rigid panel applications

Deployed CellSaver Panels
Stowed CellSaver
ABLE Patent 6,177,627 B1, and other domestic
foreign patents pending
CellSaver Panel Pair 192 x 48 Deployed Area
64
Conventional Rigid Panel Arrays

• PUMA Advanced Rigid Substrate Solar
  Array
• High performance composite honeycomb substrates
  and yoke structures
• Multiple torsion springs with viscous damper
  governed deployment
• Multiple coordination systems
• pantographic or cable/pulley coupling
• HOP actuated launch restraint/release device
• Lightweight, low cost, high reliability heritage
  design

DS1 PUMA Array
Indostar PUMA Array
GPS IIF PUMA Array
PUMA Array for MMS North American
BSAT-2 A/B PUMA Array
65
Optimized Planar Panel (ORP) Array

• Allows for a significant increase in specific
  power with a highly optimized panel, while
  utilizing conventional array subsystems
• Optimized Planar Panel (ORP) platform
• Low mass picture-frame panels sandwiched and
  snubbed between conventional panels to survive
  launch environment
• Flight-proven heritage components used for all
  other subsystems (only significant development is
  the picture frame panel)
• Conventional accordion panel deployment powered
  by spring driven hinges
• Specific power performance potential up to
  approximately 110 W/kg BOL with 27
  GaInP2/GaAs/Ge PV
• Platform also suitable for CellSaver optics which
  will provide specific power performance potential
  up to approximately 125 W/kg BOL with 27
  GaInP2/GaAs/Ge PV
• Platform also suitable for SCARLET SLA optics
  which will provide specific power performance
  potential up to approximately 180 W/kg BOL with
  27 GaInP2/GaAs/Ge PV
• Low mass achieved by
• Using reduced mass picture frame
• High photovoltaic efficiency

3m
1m
Inboard Outboard Panels Standard construction
K13C2U Graphite Composite Honeycomb Sandwich
Panels
Middle Panels Woven K13C2U Facesheet supported
by graphite honeycomb sandwich Picture Frame
all-around
SECTION A-A
Inboard Outboard Panels Standard construction
K13C2U Graphite Composite Honeycomb Sandwich
Panels
66
Next-Generation Solar Cell/Array Designs

• Multijunction Solar Cells
• Holy Grail Cells
• Thin-Film Solar Cells
• Ultra-Lightweight Arrays
• - Flexible Thin-Film
• - PowerSail Program

67
AFRL 5 Year Technical Roadmap
2000 2001 2002 2003
2004 2005
GPS, SBIRS, AEHF, Gapfiller NRO missions
35 DUST (50/50 cost share)
Crystalline Multijunction Solar Cells
29 production 31 prototype
35 DUST PH II (50/50 cost share)
32-34 (prod) 35-37 (proto
TechSat 21 PowerSail, TamArray, NRO designs
Blanket/submodule devt
11 prod blankets
Flexible- Thin-Film- Photovoltaics (FTFPV)
9 prod 12 proto
DUST (50/50 cost share)
9(prod) modules, 250W/kg array
Module/array devt 1kW LEO, 20 kW GEO,
50 kW MEO
In-house adv. FTFPV materials/device
physics/design/characterization
Quantum-Leap Concepts
holy-grail MJ- FTFPV Cell
High-temperature polymer
Spectrum compression concepts (dendrimer),
painted sol-gel FTFPV, polymer PV, direct
spectrum conversion (micro-antenna), IPC, etc...
68
30-35 Efficient 3- 4-Junction Solar Cell
DUST Program

• 5 Yr, 10M 50/50 cost share w/Spectrolab,
  Emcore, NREL
• Focused on 3-J 30 and 4-J 35 solar cell
  designs
• 4-Junction design based on AF SNL US Patent
  No. 5,944,91
• - sold to industry for 300K

• Greatest challenge
• - purity of N-source
• Best internal QE is 80
• - Jsc of 11.5 mA/cm 2
• GaInP/GaAs/GaInNAs 3-J
• cell has Voc 2.8 V

To date the program has increased commercially
available large area (27.5 cm2) 3-J solar cell
efficiency from 24.5 to 27.5, soon 28-29
69

Flexible-Crystalline Solar Cells Holy Grail
Cell (best you can do)
Objective 35 efficient, flexible, 4000 W/kg
blanket
Requirements
J1
J2
High quality MOCVD 3- or 4-junction epilayers
J3
J4
New crystalline/polymer transition layer
technology
transition layer
flexible substrate
High-temperature compatible (gt600C)
transition layer
polymer
Nanocrystal Low Temperature Recrystallization
Passivated, Ultra-large Grain Polycrystalline
Substrate-less Thin Films
70
Revolutionary Flexible-Thin-Film Photovoltaics
(FTFPV) Solar Array
State-of-Art Rigid Arrays
FTFPV Solar Array
FTFPV promising for 10X greater radiation
resistance, 3-5X lighter, 3X smaller stowed
volume, and 5X cheaper than SOTA rigid solar
arrays
71
Comparison of Solar Array Technologies(1 kW, 7
yr, 2000 km)
NEAR-TERM GPS, SBIRS Adv. EHF, Gap-Filler, classi
fied
NEXT-GEN TechSat 21 PowerSail TamArray
Crystalline III-V multijunction
Flexible-Thin-Film-PV
Flex-Thin-Film III-V MJ
Array support structure is 0.1kg/m2, FTFPV
blanket, interconnects, cabling and thermal
control included
72
FTFPV Solar Array Devt Approach
cell
FTFPV module
sub-module
1
3
2
1-2 m2
Efficiency goal 10-15
Goal 90 yield (0.5x0.5 1x1ft2)
Electrical/Mech. Integration (2-4ft2)
03-04
module delivery
05-07
03-04
PowerSail (50-100kW)
DUST - TamArray (1-20kW)
TechSat 21 (lt 1kW)
73
AFRL FTFPV Program
FTFPV Program Strategy
CIGS Cell Development
a-Si Cell Sub-Module Devt
Iowa Thin Films
ISET
DayStar
ITN
USS
Lockheed Martin DUST
Boeing DUST
FTFPV Module and Array Development Programs
PowerSail Blanket Devl
TechSat 21
Electrical Grain-Boundary and Structural Studies
Space Encapsulant Development
Radiation Studies
Array Development In-House Support
74
U.S. FTFPV Blanket Efficiencies
Substrate high-temp (gt550C)
low-temp (lt450C) properties
198g/cm2 per mil 36g/cm2 per mil
same as a-Si same as a-Si
monolithic difficult
monolithic Vendor
USS USS
AF Program SBIR PH2 SBIR PH2
w / ISET GSE Best
12(0.1ft2) 7.5 (2mil,
0.5ft2) 17 (lt1cm2)NREL 9 (0.1in2)
Efficiency 9.8(1ft2)
goal 10 (1ft2) 12
(25cm2)NREL 2 (0.1ft2)
(AM0) (5mil)
15.2
(1.1cm2)DayStar _at_425C

10.8
(24cm2)Daystar

10.2 (10cm2)ISET
painted cell
4-5 (1mil)
10ft2 11 (0.68cm2)GSE

ITF SBIR PH 7.0/8.0
(58cm2/29cm2)GSE



11-12 (4ft2 on
glass)Siemens(sputtered)
Air Force funded co-evaporation
75
a-Si Cell Sub-module Development
United Solar Systems (USS)

• Terrestrial production line for SS web
• Cells flown on Mir, space qual by Fokker
• SBIR Phase II polyimide substrate,
• web-based production process
• PowerSail blanket demo
• Sub-module efficiencies on SS
• 12, 0.1 ft2
• 9.8, 1 ft2
• on polyimide 7.5, 0.5 ft2, 10 goal

Iowa Thin Films

• Terrestrial web-based process uses a polyimide
• substrate monolithically integrated
• SBIR Phase I Develop space product from
  terrestrial
• sub-modules, projected to raise efficiency of
• terrestrial product from 4 to 9

76
USS a-Si module for Mir
United Solar Systems a-Si
77
CIGS Cell Development
International Solar Electric Technologies (ISET)

• CIGS cells manufactured in a 2 Stage Process
• 1. Deposition of Metal Precursors (Cu, In,
  Ga)
• 2. Selenization using H2Se
• Non-Vacuum Process
• Substrates Metal foil and ISET-X (proprietary
• low-weight dielectric)
• Current ISET solar cell efficiency
• 10.2 AM0 efficiency, 10 cm2
• 500 W/kg on 25 mm Mo foil
• 1/Watt projected for terrestrial cells
• BMDO SBIR piggyback funding

metal nanoparticle precursors (inks) are
deposited by a painted-on sol gel process
78
CIGS Cell Development
ITN (GSE)
DayStar

• Co-evaporated Process
• Web-based production line installed
• Thermal Cycle Testing, -196 to 100 ºC
• Thermal/ESD Protective Coating
• Development
• Welded Cell Interconnect Development
• 1 mil Stainless Steel foil substrate
• Efficiency (AM0)
• 11, 0.68cm2
• 8.5, 57cm2

• Co-evaporated Deposition
• Batch production process
• Demonstrated 1400 W/kg with 25
• mm Ti substrate
• Efficiency goal 13 AM0 for
• 20 cm2 cells
• Cell Area goal gt200 cm2
• AF 6.2 Funding

79
FTFPV Module Development Effort
(ABLE Engineering)

• Pathfinder for TechSat 21, PowerSail and
  TamArray DUST SquareRigger DUST Programs
• Phase I intensive 4 mo. 150K effort to
  identify critical FTFPV module integration issues
  
• Evaluate hardware electrical/mechanical
  characteristics and survivability for

A) a-Si and CIGS cells
B) FTFPV a-Si Demo Blanket
8-9 a-Si
7-8 CIGS
80
FTFPV Module Effort w/ABLE Engineering
At LEO Extreme Temp (-130ºC)

• Kapton Hingline Designs
• A All behind
• B Fill strip
• C Overlap
• D Overlap

A
BC
D
peeled at cold extremes
puckers at cold extremes
didnt warp!
Tape Peel-Test (grid, a-Si CIGS ahhession)
25ºC
Hingline Pull-Test (silicone adhessive)
GEO Extreme _at_ -180ºC
D
D
- metal warps - hingeline stays flat
- samples pulled _at_ gt25 lbs/in - blanket tension
lt 0.05 lbs/in
- looks good after LEO temps. - peeling after
GEO temps.
flat

• welding cells into module
• electrical characterization
• thermal cycling (-180 to 120ºC)

Identified low temperature cycling as a
challenge for some cell designs
81
FTFPV Space Qualification Program (via
MIL-STD-1540C)
FTFPV Space Qualification
Degradation vs Proton Exposure

• HIGH TEMPERATURE PERFORMANCE (TEMP
  COEFFICIENTS)
• ON-ORBIT TEMPERATURE MODELING (AFRL)
• TEMPERATURE COEFFICIENCTS _at_ 25-100C (AFRL)
• RADIATION HARDNESS (NRL and AEROSPACE CORP.)
• ELECTRONS (0.5 - 2 MEV UPTO 1E15 /cm2)
• PROTONS (50KEV - 10 MEV UPTO 1E13/cm2)
• ANNEALING AT 70-100C
• THERMAL CYCLING (EMCORE PV)
• -110 TO 150C AT 80 C/MIN
• CONTACTS, INTERCON.,
• ENCAPSULANTS DELAMINATION
• ELEC/MECH INTERCONNECT STRENGTH/STRESS
• DUE TO THERMAL
• AF DUST, SBIR PHII w/ABLE ENGINEERING
• MECHANICAL DURABILITY (LAUNCH DEPLOYMENT)
• AF DUST, SBIR PHII w/ABLE ENGINEERING

a-Si
Displacement Damage Dose (MeV/g)
CIGS
No show-stoppers have surfaced yet for FTFPV
cells and blankets
82
FTFPV Array Issues

• Significant technical progress achieved,
  significant govt and industry interest/
• investment exist, potential pay-off is
  significant/ enabling for many applications
• However, issues remain
• ? thin-film solar cells efficiency 1/3
  efficiency crystalline solar cells
• ? FTFPV array area 3x of 3-J solar cell
  array area
• ? increased propellant for reaction gyro
  system desaturation and drag makeup
• ? field-of-view issues
• ? low first fundamental frequency
• ? propulsion system thrust level limit
• 1st order analysis on aerodynamic drag/solar
  pressure effects
• by Dave Hoffman, NASA Glenn (next charts)
• Present FTFPV development drivers
• ? DoD - high power
• ? ComSats cost

83
Thin-Film Solar Array Earth-Orbit Mission Study
Dave Hoffman
Dave Hoffman NASA Glenn david.j.hoffman_at_
grc.nasa.gov [phone]
Glenn Research Center
84
Thin-Film Solar Array Earth-Orbit Mission Study
Dave Hoffman
Dave Hoffman NASA Glenn david.j.hoffman_at_
grc.nasa.gov [phone]
Glenn Research Center
65 W/kg
325 W/kg
85
FTFPV Dual-Use-Science Technology (DUST)
Program

• Programmatics
• 3yr/8M (50/50 govt/industry)
• Primes LockMart Denver
• and Boeing
• Subs LockMart Sunnyvale,
• ABLE Engineering, ISET ITN

Define basic module design
3
2
Space qualify module, module-to-support
structure Integration and test
1-3ft

• Qual. Testing
• radiation
• temp cycle
• vibration
• O UV
• high voltage

Identify preferred solar array support structure
design
1-3ft
1
modules
Module-to-structure interconnects
tabs
kapton
Representative support structure
Phase IV Option Flight Demo ( 5-10kW)
4
Encapsulant development High
emissivity, high dielectric constant, dense,
transparent, good anti-reflection properties

86
Lockheed Martin FTFPV DUST Program

• Thin Film Photovoltaics for Next Generation
  Solar Arrays
• 4.3M over 3 years, 50 - 50 cost share between
  
• industry government
• LM Astronautics, USS, ISET subcontractors
• 5 kW, 10 kW, 20 kW FTFPV module technology
• Specific Objectives Develop demonstrate
• stabilized 10-15efficient FTFPV sub-modules
• sub-module and module electrical architectures
• including bypass and blocking diode technology
• module strength and structural support
  requirements

Optimize cells FTFPV array design
Optimize sub-modules fabricate modules
PV down select space qualification
Phase I
Phase II
Phase III
Kickoff 7/01
3/02
3/03
3/04
Program Timeline
87
Boeing FTFPV DUST Program

• Ultra-Lightweight Thin-Film Photovoltaics Module
  
• Technology
• 4M over 3 years, 50 - 50 cost share between
• industry government
• Able Engineering, USS, ITN subcontractors
• 1, 20, 50 kW FTFPV module technology
• Specific Objectives Develop demonstrate
• stabilized 10-15efficient FTFPV sub-modules
• sub-module and module electrical architectures
  including bypass and blocking diode technology
• module strength and structural support
  requirements

88
AFRL PowerSail ProgramAFRL Inhouse/Able
Engineering

• PowerSail is an extremely large area
  deployable solar array system (gt50 to 100 kW
  BOL)
• Free-flyer with integral ACS, PMAD, Control Bus
• Power transferred to host S/C through attachment
  system - Thin film or Crystalline PV
• Gossamer structural areal density (lt 0.2 kg/m2),
  and stowed volume (gt25 kW/m3)
• Technology is enabling for advanced Air Force
  applications

89
PowerSail Program
Program Manager Nick Hague

• Team ARRL, other govt agencies, Able, Cal
  Tech, Texas AM, U. of Colorado
• Leverages Ables SquareRigger array design and
  AFRL FTFPV devt programs
• 20kW flight demo in 05-06
• Funding 022M, 034M, 047M, 0512M,
  0610M, 075M

Square Rigger Array Design/Deployment
90
PowerSail Performance Comparison
91
Mechanical Deployment System Demo
92
PowerSail One-Bay Hardware Deployment Demo
93
PowerSail One-Bay Hardware Deployment Demo
94
PowerSail One-Bay Hardware Deployment Demo
95
PowerSail One-Bay Hardware Deployment Demo
96
PowerSail One-Bay Hardware Deployment Demo
97
PowerSail One-Bay Hardware Deployment Demo
98
PowerSail One-Bay Hardware Deployment Demo
99
PowerSail One-Bay Hardware Deployment Demo
100
PowerSail One-Bay Hardware Deployment Demo
101
Summary

• High efficiency multijunction solar cells offer
  economic advantages
• at spacecraft systems level highest
  efficiency cells available utilized
• on most DoD and commercial ComSats
• Near-term (1-3yrs)
• 28-30 efficient 3-junction cells commercially
  available
• ? 100-150 W/kg solar array
• Far-Term (3-10yrs)
• - 33-36 efficient 4-junction cells
  commercially available
• ? 150-200 W/Kg solar array
• - 12-20 efficient thin-film space solar
  cells commercially available
• ? 300-500 W/Kg solar array
• ? 50-100 kW array