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i

NASA/TM—2011–216476

Advanced X-Ray Timing Array Mission: 

Conceptual Spacecraft Design Study

November 2011

National Aeronautics and

Space Administration

Marshall Space Flight Center • Huntsville, Alabama 35812

R.C. Hopkins, L. Johnson, H.D. Thomas, and C.A. Wilson-Hodge

Marshall Space Flight Center, Huntsville, Alabama

M. Baysinger and C.D. Maples

Qualis Corporation, Huntsville, Alabama

P.D. Capizzo

Raytheon Company, Huntsville, Alabama

L.L. Fabisinski

International Space Systems, Inc., Huntsville, Alabama

L. Hornsby and K.S. Thompson

Jacobs ESTS Group, Huntsville, Alabama

J.H. Miernik

ERC, Inc./Jacobs ESTS Group, Huntsville, Alabama
ii

Available from:

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076 –1320

443–757–5802

This report is also available in electronic form at

<https://www2.sti.nasa.gov/login/wt/>
iii

TABLE OF CONTENTS

1. INTRODUCTION ............................................................................................................. 1

2. SCIENCE MISSION SUMMARY .................................................................................... 2

3. SCIENCE INSTRUMENTS .............................................................................................. 3

4. MISSION AND SPACECRAFT REQUIREMENTS ....................................................... 5

5. MISSION ANALYSIS ........................................................................................................ 6

5.1 Initial Spacecraft Mass Estimate .................................................................................. 6

5.2 Launch Vehicle Selection and Performance .................................................................. 7

5.3 Orbital Lifetime ............................................................................................................ 9

5.4 End of Life Disposal .................................................................................................... 9

5.5 Miscellaneous Tasks ..................................................................................................... 11

6. SPACECRAFT DESIGN FOR THE TAURUS II LAUNCH VEHICLE ......................... 12

6.1 Configuration ............................................................................................................... 12

6.2 Mass Properties ............................................................................................................ 13

6.3 Propulsion .................................................................................................................... 14

6.4 Communications .......................................................................................................... 15

6.5 Avionics, Guidance, Navigation, and Control ............................................................... 16

6.6 Power ............................................................................................................................ 17

6.7 Structures ..................................................................................................................... 19

6.8 Thermal ........................................................................................................................ 21

7. PRELIMINARY SPACECRAFT DESIGN 

FOR THE FALCON 9 LAUNCH VEHICLE ................................................................... 27

7.1 Configuration ............................................................................................................... 27

7.2 Mass Properties ............................................................................................................ 28

7.3 Propulsion .................................................................................................................... 29

7.4 Communications, Avionics, Guidance, Navigation, and Control .................................. 29

7.5 Power ............................................................................................................................ 30

7.6 Structures ..................................................................................................................... 31

7.7 Thermal ........................................................................................................................ 32

8. RISK ANALYSIS ............................................................................................................... 33
iv

TABLE OF CONTENTS (Continued)

 9. ANIMATION ................................................................................................................... 35

10. CONCLUSIONS .............................................................................................................. 36

REFERENCES ....................................................................................................................... 37
v

LIST OF FIGURES

 1. Cutaway rendering of a LATA super module ............................................................ 3

 2. Rendering of (a) an SM camera and (b) a cluster of seven cameras .......................... 4

 3. Plot of spacecraft payload mass as a function of dry bus mass 

from the Rapid II catalog .......................................................................................... 6

 4. Conceptual layout of the 20 LATA super modules (each 60×60 cm) 

in the Minotaur IV and Taurus II shrouds ................................................................ 8

 5. Plot of launch vehicle performance for several vehicles. Final inclination could 

be traded for spacecraft mass if needed for the AXTAR mission. All launches 

are from Cape Canaveral Air Force station (CCAFS) unless otherwise noted ........... 8

 6. AXTAR orbital lifetime trade ................................................................................... 9

 7. Ratio of finite burn to impulsive ∆Vs as a function of thrust acceleration ................. 10

 8. AXTAR configuration for the Taurus II launch vehicle option ................................. 12

 9. AXTAR propulsion system schematic ....................................................................... 14

10. Solar-based power system ......................................................................................... 18

11. FEMAP Taurus II model deformed side view showing rows of Sun shades and 

exaggerated deformations .......................................................................................... 21

12. Thermal model geometry for the Taurus II configuration ......................................... 22

13. Orientation of spacecraft showing the (a) cold case and (b) hot case used 

in the thermal analysis ............................................................................................... 22

14. Analysis results for the Taurus II configuration: Spacecraft bus (cold case) .............. 24

15. Analysis results for the Taurus II configuration: LATA structure (cold case)............. 24

16. Analysis results for the Taurus II configuration: Spacecraft bus (hot case) ................ 25

17. Analysis results for the Taurus II configuration: LATA structure (hot case) .............. 25
vi

LIST OF FIGURES (Continued)

18. Configuration of the AXTAR spacecraft: Falcon 9 option ....................................... 27

19. FEMAP model used for the Falcon 9: (a) Side view and (b) top view ....................... 31
vii

LIST OF TABLES

 1. Spacecraft and mission requirements and instrument data ........................................ 5

 2. Initial payload mass estimate for AXTAR ................................................................ 7

 3. Initial estimate for the spacecraft mass ...................................................................... 7

 4. Deorbit propellant requirements for the Taurus II option ......................................... 10

 5. Deorbit propellant requirements for the Falcon 9 option .......................................... 10

 6. AXTAR mass summary for the Taurus II configuration ........................................... 13

 7. Propulsion system mass summary for the Taurus II option ....................................... 15

 8. Avionics mass results for the Taurus II configuration ................................................ 17

 9. Power budget for the AXTAR spacecraft .................................................................. 17

10. Power system mass summary ..................................................................................... 18

11. AXTAR structures mass summary for the Taurus II configuration ........................... 21

12. Thermal model surface optical properties ................................................................. 23

13. Heat dissipation, operating, and predicted temperatures 

for the Taurus II configuration .................................................................................. 23

14. Thermal subsystem mass summary for the Taurus II configuration .......................... 26

15. AXTAR mass summary for the Falcon 9 configuration ............................................ 28

16. Propulsion system mass summary for the Falcon 9 option ........................................ 29

17. COMM, avionics, and GN&C mass summary for the Falcon 9 configuration ........... 30

18. Power budget for AXTAR for the Falcon 9 configuration ......................................... 30

19. Power system mass summary for the Falcon 9 configuration ..................................... 30
viii

LIST OF TABLES (Continued)

20. Structural mass summary for the AXTAR spacecraft 

for the Falcon 9 configuration ................................................................................... 31

21. Thermal system mass summary for the AXTAR spacecraft 

for the Falcon 9 configuration ................................................................................... 32

22. AXTAR heat dissipation and operating temperatures for the Falcon 9 ..................... 32

23. Summary data for the two spacecraft designs ............................................................ 36
ix

LIST OF ACRONYMS AND SYMBOLS

ACS attitude control system

AIAA American Institute of Aeronautics and Astronautics

AITECH AITech International

Al aluminum

ATK Alliant Techsystems, Inc.

AXTAR Advanced X-ray Timing Array

BAT Burst Alert Telescope

BH black hole

CCAFS Cape Canaveral Air Force Station

CG center of gravity

CMG control moment gyros

COMM communication

COTS commercial off the shelf 

Cu copper

DAS Debris Assessment Software

FCI Fluid Components International

FEMAP Finite Element Modeling Analysis and Postprocessing

GaAs gallium arsenide

GSFC Goddard Space Flight Center

IDS Instrument Data System
x

LIST OF ACRONYMS AND SYMBOLS (Continued)

IMU Inertial Measurement Unit

ITJ improved triple junction

LATA Large Area Timing Array

Li lithium

MEL master equipment list

MGA mass growth allowance

MIDEX medium-class Explorer

MLI multilayer insulation

MOS margin of safety

MPS main propulsion system

MSFC Marshall Space Flight Center

NS neutron star

ORS Sat-1 operationally responsive space satellite 1

pyro pyrotechnic

RCS reaction control system

RSDO Rapid Spacecraft Development Office

RW reaction wheel

RXTE Rossi X-Ray Timing Explorer

SAA South Atlantic Anomaly

Si silicon
xi

LIST OF ACRONYMS AND SYMBOLS (Continued)

SM Sky Monitor

STK Satellite Took Kit

TDRSS Tracking and Data Relay Satellite System

TM Technical Memorandum

TRL Technology Readiness Level

UV ultraviolet

VME Versa Module Europa

ZnO zinc oxide
xii

NOMENCLATURE

α absorptivity

g acceleration due to gravity at standard sea level

ΔV delta velocity (change in velocity of an orbital maneuver)

m0 initial mass

ε emissitivity

ε* effective emissitivity
1

TECHNICAL MEMORANDUM

ADVANCED X-RAY TIMING ARRAY MISSION: 

CONCEPTUAL SPACECRAFT DESIGN STUDY

1. INTRODUCTION

In the spring of 2010, the Advanced Concepts Office of the NASA Marshall Space Flight 

Center (MSFC) completed a conceptual spacecraft design to meet vehicle requirements for the 

Advanced X-Ray Timing Array (AXTAR) science mission. The goal was to design a spacecraft 

that provides power and data handling for the instruments, communication (COMM), pointing, 

station keeping, momentum unloading, thermal control, and end-of-life disposal. Since the science 

team was interested in both a minimum-mass design and one that carried the maximum possible 

number of instruments, two separate conceptual designs were created. The approach was to first 

complete a conceptual spacecraft design that could accommodate a minimum-required instrument 

suite as specified by the science team and select a possible launch vehicle (or family of vehicles), 

and then complete at least one iteration of a design exercise in which the maximum number of 

science instruments were accommodated by a spacecraft launched on the Falcon 9 launch vehicle. 

The smaller configuration, which was the primary product of the study, is a closed design, while the 

more massive configuration was not iterated to closure due to time constraints. Nevertheless, the 

larger configuration is representative of a science package designed to maximize the capability 

of a Falcon 9 or similar launch vehicle.

An introduction to the science mission, the science instruments, and the details of the spacecraft requirements and design are included in the sections of this Technical Memorandum (TM).
2

2. SCIENCE MISSION SUMMARY

This TM presents a brief summary of AXTAR’s science mission and requirements. See 

reference 1 for more details.

AXTAR’s primary science goals focus on extracting fundamental physics from neutron stars 

(NSs) and stellar mass black holes (BHs) (the end products of stellar evolution for massive stars). 

An NS is composed of ultradense matter with the mass of the Sun compressed into a star the size 

of a city. Stellar mass BHs are thought to contain the mass of several to tens of suns compressed 

into a singularity (a single point) surrounded by an event horizon, where the escape velocity from 

the BH equals the speed of light. Orbital periods near the BH event horizon are on millisecond 

timescales and are set by the BH’s mass, angular momentum, and the laws of relativistic gravity. 

Maximum rotation periods of NSs are also on millisecond timescales and are set by the equation 

of state, the relationship between the star’s mass and radius, and of ultradense matter in its interior.

With its large area and high time resolution, AXTAR’s Large Area Timing Array (LATA) 

will address the following key science questions raised in the 2010 Astronomy and Astrophysics 

Decadal Survey: 

(1) How do BHs work and influence their surroundings?

(2) How do rotation and magnetic fields affect stars?

(3) What controls the masses, spins, and radii of compact stellar remnants?

The Sky Monitor (SM) instrument on-board AXTAR will serve as a trigger for LATA 

observations and will monitor hundreds of x-ray sources for a wide variety of primary science 

investigations. Partnering with wide-angle astronomy observatories at other wavelengths, the SM 

will support time-domain astronomy including surveying the sky at many wavelengths (stressed 

as important science in the 2010 Astronomy and Astrophysics Decadal Survey).

Reference 1 gives the science objectives that set AXTAR’s mission requirements in detail. 

A few key points are summarized in this TM. AXTAR’s required effective area is set by two key 

science objectives: (1) Achieving a 5–10% NS radius measurement and (2) achieving a 0.1% rms 

fractional amplitude detection threshold for high-frequency, quasi-periodic oscillations in BHs. Its 

targets will be bright galactic sources, requiring that the detector and data system handle high data 

rates without impacts from dead time, pileup, or data losses. Many of the target observations need 

to be triggered when a source is in a particular state, and a sensitive SM with maximal coverage 

of the accessible sky is required because x-ray transients may occur anywhere in the sky. The Rossi 

X-Ray Timing Explorer (RXTE) has shown that AXTAR’s targeted events are often short lived, 

requiring the capability to reschedule in a matter of hours and slew to new coordinates 

expeditiously.
3

3. SCIENCE INSTRUMENTS

While some basic information on the science instruments is included for completeness, 

this TM focuses on the spacecraft design. Refer to reference 1 for details regarding the science 

instruments.

The two science instruments for AXTAR are the LATA and the SM (which is a collection 

of cameras). The function of the SM is to view as much of the sky as possible, detect x-ray 

sources, and trigger the spacecraft to slew and point the LATA at the x-ray source for detailed 

observations. The function of the Instrument Data System (IDS) is to process the data from the 

science instruments. A short summary of the instrument parameters is found in section 4.

The LATA is the primary AXTAR science instrument and is composed of several coaligned 

super modules (see fig. 1). From top to bottom, the components in figure 1 are as follows: The collimator, a light shield, silicon (Si) detectors, an interposer board, a digital board, and a mounting 

plate that provides support and shielding. The 5×5 array of Si detectors (each 10×10 cm) results 

in a super module length and width of approximately 60 cm due to space between the detectors and 

border material. With the collimator installed, the height of each super module is approximately 

20 cm. As stated above, all super modules are coaligned to within 1 arc-min, although they do not 

have to be coplanar. The science team desired a spacecraft configuration that contained at least 20 

super modules.

Collimator

Light Shield

Si Detectors

Interposer Board

Digital Board

Mounting Plate

60 cm

20 cm

60 cm

Figure 1. Cutaway rendering of a LATA super module.
4

The SM observes a broad portion of the sky and provides the data necessary to repoint the 

LATA at an active x-ray source. The SM is composed of a maximum of 32 cameras, each pointing 

to a different part of the sky except for a small area of overlap. Figure 2 shows an illustration of a 

single SM camera and a cluster of seven cameras. Each SM camera contains a four-element detector plane (see fig. 2(a)) consisting of the same Si pixel detectors used for the LATA and is topped 

with a two-dimensional coded mask. The base of each camera is 25×25 cm, and the top is 30-cm 

square, with a total height of 25 cm. The spacecraft carries a maximum of 32 SM cameras in an 

arrangement similar to a soccer ball. Each camera looks in a different direction, resulting in total 

sky coverage.

The SM cameras do not have to be collocated on the spacecraft, but can be placed on various locations for convenience and packaging. The science team desired a minimum of seven SM 

cameras, with the central field of view of the cluster being coaligned with the LATA. If another 

cluster of seven could be added, the next priority was to point the cluster in the opposite direction 

of the first. This is because the x-ray sources predominately lay along the galactic plane. If room 

and power were available, additional SM cameras could then be placed on the spacecraft, up to 

a maximum of 32.

(a) (b)

Coded Mask

Detector 

Plate

30 cm 30 cm

25 cm

25 cm 25 cm

Figure 2. Rendering of (a) an SM camera and (b) a cluster of seven cameras.
5

4. MISSION AND SPACECRAFT REQUIREMENTS

The top-level requirements for the AXTAR design study are listed in table 1. Major parameters for the science instruments (the LATA, SM, and IDS) are also listed for completeness. The IDS 

is not actually an instrument, but it is included as part of the science payload because it processes 

the data from the SM and LATA. The spacecraft structure and subsystems were designed to meet 

these requirements and included a 30% mass growth allowance (MGA) margin on both dry mass 

and power for the spacecraft and instruments.

Table 1. Spacecraft and mission requirements and instrument data.

Parameter Required (Desired or Bounding) Value

Spacecraft and Mission

Anticipated launch year

Orbit altitude

Orbit inclination

Spacecraft lifetime

Consumables

Orbit lifetime

Pointing accuracy

Pointing knowledge

Maximum slew rate

2019 (used as a guide for orbit lifetime estimates)

Approximately 600 km (study output), circular

28.5° or less (as low as possible)

3 yr

5 yr

10 yr

1 arc-min (or better)

5 arc-s (or better)

180° in 30 min

SM Camera (each)

Mass

Power

Dimensions

Quantity

Thermal requirement

Alignment

2 kg + 2 kg per telemetry hub

4 W + 9 W per telemetry hub

30 × 30 × 25 cm

7 minimum (32 maximum)

–400 °C to +100 °C (detector plane)

32 faces and vertices of a dodecahedron.

LATA Super Module (each)

Mass

Power

Dimensions

Quantity

Thermal requirement

Alignment

30 kg

30 W

60 × 60 × 20 cm

20 (as many as possible)

–400 °C to + 100 °C (detector plane)

All coaligned within 1 arc-min

IDS

Mass

Power

Dimensions

Quantity

20 kg

40 W

30 × 30 × 20 cm

1

Contingency Philosophy

Mass

Power

30% for spacecraft subsystems and instruments.

30% for spacecraft and instruments.
6

5. MISSION ANALYSIS

The mission analysis tasks included generating an initial spacecraft estimate, gathering 

launch vehicle performance data and recommending one or more vehicles to define the dynamic 

envelope, estimating the orbit lifetime, determining the amount of propellant required for a controlled deorbit, and investigating ground station contact times for various orbit inclinations. Several tools (e.g., NASA Debris Assessment Software (DAS), Analytical Graphics’ Satellite Tool Kit 

(STK), and in-house tools) were used to complete these assessments.

5.1 Initial Spacecraft Mass Estimate

To determine an initial spacecraft mass estimate, the team looked at historical spacecraft 

bus data from the Goddard Space Flight Center (GSFC) Rapid Spacecraft Development Office 

(RSDO) catalog, version 2, known as Rapid II. Rapid II was chosen since the new Rapid III 

catalog had not yet been released. Plotting the payload mass capability of the buses versus their 

dry masses resulted in what appeared to be two groups of spacecraft (see fig. 3). The team briefly 

attempted to determine the reason for the two groups, which consisted of a ‘light’ group where 

the payload outweighed the dry bus and a ‘heavy’ group where the dry bus outweighed the payload. However, after looking at power, propulsion, and other subsystem data, the team could not 

find any trends that could explain the two apparent groups. Therefore, the mission analysis team 

decided to use the average of the light and heavy estimates as the basis for the initial dry bus mass 

estimate. In order to determine the bus mass estimate, the payload (science instrument) mass must 

be found.

1,000

900

800

700

600

500

400

300

200

100

0

0 100 200 300 400 500 600 700 800 900 1,000

y=1.77x

y=0.676x

Bus Mass, Dry (kg)

Payload Mass (kg)

Heavy Group

Light Group

Figure 3. Plot of spacecraft payload mass as a function of dry bus mass 

from the Rapid II catalog.
7

The payload consists of two science instrument suites, the LATA and a collection of several 

SM cameras. The LATA is composed of 60-×60-cm super modules, each with a mass of 30 kg. 

The SM cameras have a mass of only 2 kg each. Since the science team desired a minimum of 20 

LATA super modules, 7 SM cameras, and a 20-kg IDS, the estimated total payload mass was 634 

kg (see table 2). Given this mass and the curve fits of figure 3 and adding preliminary propellant 

capabilities for deorbit, the team arrived at an initial total observatory mass (spacecraft, science 

instruments, and propellant) of approximately 2,000 kg (see table 3). The final observatory mass 

resulting from the design study was almost equal to that predicted by the ‘heavy’ spacecraft bus 

curve. Nevertheless, the initial estimate was an effective starting point for analysts to begin sizing 

subsystems.

Table 2. Initial payload mass estimate for AXTAR.

Instrument Mass Each Qty Mass Total

LATA

SM

IDS

30

2

20

20

7

1

600

14

20

Total instrument mass 634

Table 3. Initial estimate for the spacecraft mass.

Component

Mass 

(Low Estimate)

Mass 

(High Estimate)

Payload

S/C dry mass

Contingency (30%)

Subtotal dry mass

Propellant

634

450

325

1,409

90

634

1,200

550

2,384

150

Approximate total mass 1,500 2,500

Split the difference between the low and high estimates ≈2,000 kg

5.2 Launch Vehicle Selection and Performance

Given the basic dimensions of the LATA super modules (60×60 cm) and the desire to have 

a minimum of 20 LATA super modules, the team created a preliminary layout of the LATA to 

compare its dimensions with several launch vehicle shrouds, two of which are shown in figure 4. 

Since it appeared that the LATA in the Minotaur IV shroud left too little room for the spacecraft 

(and was a tight fit by itself) and virtually no mass margin, the team decided to baseline the Taurus II as the launch vehicle. Given this decision, the smaller configuration with 20 LATA super 

modules could be launched on either a Taurus II or Falcon 9 (or larger) vehicle, while the larger 

configuration with the maximum number of super modules would require a Falcon 9 or larger 

vehicle.
8

Minotaur IV

Taurus II

20 LATA Modules, 

60 × 60 cm per Module.

Figure 4. Conceptual layout of the 20 LATA super modules (each 60×60 cm) 

in the Minotaur IV and Taurus II shrouds.

At the completion of the spacecraft design cycle, this decision was revisited to make sure 

that the spacecraft mass and volume were still within the limits of the selected launch vehicles. 

Section 6.1 provides more details.

Given a rough idea of the launch vehicle size needed, the team gathered payload mass 

capability data for a representative cross section of vehicles as a function of final inclination. 

A plot of several vehicle performance curves is shown in figure 5. Even though it is desirable 

to minimize passage through the South Atlantic Anomaly (SAA) by going to as low an inclination 

as possible, the science team was willing to trade inclination for more LATA super modules. Therefore, performance to various inclinations was needed. The Atlas V 401 performance was included 

for comparison purposes, and the Taurus II enhanced performance is very preliminary.

7,000

6,500

6,000

5,500

5,000

4,500

4,000

3,500

3,000

2,500

2,000

1,500

1,000

500

0

Payload Mass (kg)

0 5 10 15 20 25 30

Final Orbital Inclination (deg)

Atlas V 401

Falcon 9 

from 

Kwajalein

Launch vehicle 

performance to 600-km 

circular orbit versus 

final orbit inclination.

Falcon 9

Taurus II Enhanced

Taurus II

Figure 5. Plot of launch vehicle performance for several vehicles. Final inclination could 

be traded for spacecraft mass if needed for the AXTAR mission. All launches 

are from Cape Canaveral Air Force station (CCAFS) unless otherwise noted.
9

5.3 Orbital Lifetime

The NASA DAS (version 2.0.1) is used to determine the orbital lifetimes at several area-tomass values and initial circular orbit altitudes. Figure 6 shows the results of the trade analysis. The 

desire is to find the lowest starting altitude with an orbital lifetime of at least 10 years but no more 

than 25 years. The AXTAR area-to-mass is expected to vary between approximately 0.007 and 

0.011 m2/kg (bounds the low and high values for both the Taurus II and Falcon 9 configurations). 

Taking this into consideration, an initial circular orbital altitude of 585 km (preliminary) 

is selected.

25

20

15

10

5

0

Orbit Lifetime (yr)

500 525 550 575 600 625 650

Initial Circular Orbit Altitude (km)

0.006 m2/kg

0.0075 m2/kg

0.01 m2/kg

0.0125 m2/kg

0.0150 m2/kg

0.0175 m2/kg

0.02 m2/kg

Figure 6. AXTAR orbital lifetime trade.

5.4 End of Life Disposal

The same strategy planned for deorbiting the Compton Gamma Ray Observatory is used 

for the AXTAR end-of-life disposal.2 The deorbit is performed over a number of maneuvers that 

lower the perigee from the initial 585-km circular orbit. The final burn begins with an apogee of 

585 km and a perigee of 150 km and targets a controlled reentry with a flight path angle of –1.2° 

at an altitude of 60 km. The number of maneuvers (five in this case) is chosen such that the ΔVs are 

roughly equal. The first four burns have an impulsive ΔV of 30.74 m/s and the final ΔV is 37.89 m/s.

Given the impulsive ΔVs above, a finite-burn analysis is performed to determine the total 

ΔV for each maneuver, including gravity losses, for various thrust levels. The ratio of finite burn 

to impulsive ΔVs as a function of thrust acceleration for the final deorbit maneuver is shown in 

figure 7. The first four maneuvers are close in magnitude and start from the same altitude as the 

final burn; therefore, they are assumed to have the same variation. These results are used to 

compute the total deorbit propellant requirements at desired thrust levels.
10

1.40

1.35

1.30

1.25

1.20

1.15

1.10

1.05

Finite-Burn 

1.00

∆

V/Impulsive 

∆

V

Initial Thrust Acceleration (Earth g)

0 0.005 0.010 0.015 0.020 0.025 0.030

Vfb

Vimp

= 1

a + b aT (0)

ge

c

a = 1.0000

b = –3.1524×10–6

c = –1.9630

Data

Curve-fit

Figure 7. Ratio of finite burn to impulsive ∆Vs as a function of thrust acceleration.

Applying the above maneuver strategy and finite-burn results, the total maneuver propellant 

is computed for the Taurus II and Falcon 9 spacecraft configurations. The Taurus II propellants 

are summarized in table 4, which gives a total required maneuver propellant of 194.8 kg (includes 

the main and attitude control amounts). Note that the thrust level drops as the pressure decreases 

in the tank (from 420 to 110 psia, see section 6.3). The maneuver propellants are computed using 

the minimum specific impulses of the main and attitude control thrusters (i.e., 223 s and 229 s, 

respectively). Due to the preliminary nature of this analysis, the attitude control propellant is computed assuming that 5 lbf of thrust operates over the maneuver times. Table 5 gives the results for 

the Falcon 9 option, which requires 363.48 kg of propellant. Off-the-shelf tanks are used for both 

spacecraft options. These tanks hold maneuver propellant amounts of 207.94 kg and 405.25 kg for 

the Taurus II and Falcon 9 configurations, respectively (see sections 6.3 and 7.3).

Table 4. Deorbit propellant requirements for the Taurus II option.

Event

m0

(kg)

ΔV_ideal 

(m/s)

MPS Thrust 

(lbf)

ΔV_real 

(m/s)

Maneuver 

Time

(s)

Manuever

Propellent

(kg)

Deorbit Maneuver 1

Deorbit Maneuver 2

Deorbit Maneuver 3

Deorbit Maneuver 4

Deorbit Maneuver 5

2,650.30

2,613.26

2,576.73

2,540.67

2,505.05

30.74

30.74

30.74

30.74

37.89

257.40

216.05

174.70

133.35

92.00

30.77

30.79

30.82

30.88

38.24

70.7

83.1

101.5

131.3

232.1

37.73

37.35

37.26

36.92

45.73

Table 5. Deorbit propellant requirements for the Falcon 9 option.

Event

m0

(kg)

ΔV_ideal 

(m/s)

MPS Thrust 

(lbf)

ΔV_real 

(m/s)

Manuever 

Time

(s)

Manuever

Propellent

(kg)

Deorbit Maneuver 1

Deorbit Maneuver 2

Deorbit Maneuver 3

Deorbit Maneuver 4

Deorbit Maneuver 5

4894.73

4826.11

4758.33

4691.28

4624.76

30.74

30.74

30.74

30.74

37.89

257.40

216.05

174.70

133.35

92.00

30.88

30.93

31.03

31.23

39.15

131.1

154.2

188.7

245.2

438.5

69.92

69.31

68.91

68.95

86.39
11

5.5 Miscellaneous Tasks

In addition to the above analyses, analysts also computed the beta angle history for 5-yr, 

eclipse durations over a 1-year period, estimated passage durations through the SAA, and ground 

station contact times for the first 2 months. The baseline launch date for the calculations was 

June 1, 2017. The 2017 launch date was the initial value used for the study. Although this date was 

later revised to 2019, there was insufficient time to revise the lifetime analysis. However, minimal 

impact is expected as the altitude can be raised to negate any shortening of the lifetime due to 

the solar cycle. The baseline orbits for these assessments were circular (5° and 28.5° inclinations) 

with an altitude of 585 km. These results were provided as inputs to thermal, power, and COMM 

experts during the spacecraft design process.
12

6. SPACECRAFT DESIGN FOR THE TAURUS II LAUNCH VEHICLE

This section details the spacecraft design for the smaller of the two AXTAR configurations with each spacecraft subsystem being described in an appropriate subsection. Carrying the 

minimum 20 LATA super modules, this design is compatible with the Taurus II payload mass and 

volume capabilities. 

6.1 Configuration

The driving factor in the spacecraft’s configuration was the launch vehicle size, which limited the number of the primary science instruments (LATA super modules). The LATA coplanar 

and pointing requirements drove the design layout to be divided into two main sections: A spacecraft system bus and a science bus (see fig. 8). The science bus/LATA array was set forward to 

provide viewing separation from the spacecraft’s solar arrays and other systems, and the science 

bus structure was designed to allow the LATA super modules to be nestled inside the primary 

structure to provide additional shielding. In addition to the primary vehicle structure, a separate 

Sun-shade secondary structure was placed above the LATA to help meet the Sun avoidance-angle 

plane requirements. This gave the required shielding height necessary for the rows of LATA super 

modules.

SM Cluster (7)

SM Cluster (6)

SM Cluster (4)

SM Cluster (4)

Solar Array (2×)

20 LATAs (4×5)

Figure 8. AXTAR configuration for the Taurus II launch vehicle option.

The other design factor was the number and placement of the SM cameras. A 32-face 

dodecahedron would allow full-sky coverage, but 27 were chosen for the baseline configuration due 

to volume and operational constraints. These were broken up into several clusters and placed on 
13

the spacecraft to optimize view angles while minimizing size. These camera clusters were fairly low 

mass; therefore, no additional provisions were needed to accommodate them.

The spacecraft system bus is located aft near the launch vehicle interface to carry launch 

vehicle loads and to allow for the spacecraft’s systems to be enclosed in a more compact volume. 

The central core of the spacecraft bus houses the propulsion tanks and the momentum wheels in 

a pyramid configuration. The outer bays/walls allow for mounting of the other avionics and spacecraft components. The design has continuous load paths and structures that allow for minimum 

mass to be obtained. Locating the spacecraft bus aft allows for desirable center of gravity (CG) 

characteristics. The overall configuration is conservative and does allow room for component 

growth and for extra subsystem components to be added that were not analyzed in this study. 

While based on the Taurus II launch vehicle, the design could be used on comparable (not smaller) 

size launch vehicles. If a larger launch vehicle were selected, a different overall configuration and 

layout would need to be studied to achieve the optimal configuration. 

6.2 Mass Properties

A collaborative engineering environment with discipline subsystem experts was used to 

size spacecraft subsystems and propellant loads. Two design concepts were analyzed: One to be 

launched on a Taurus II or similar vehicle and one to be launched on a Falcon 9 or similar vehicle. 

Mass properties for both Falcon 9 and the Taurus II designs used 30% as the accepted contingency. 

Science instrument masses also included a 30% contingency for both designs.

The mass breakdown for the overall vehicle, spacecraft subsystems, and science instruments 

for the Taurus II design is shown in table 6. The Taurus II design resulted in a total vehicle gross 

mass of 2,690 kg. Gross mass is the total of vehicle dry mass, inert mass, and propellant. Dry mass 

is defined as spacecraft subsystems mass minus the useable propellant, propellant residuals (see 

item 7.0 in table 6), and science instruments. The dry mass total, including the 30% contingency, 

resulted in a mass of 1,567 kg. Inert mass includes propellant residuals and science instruments. 

The inert mass totaled 916 kg for the Taurus II design. The total mass less propellant (dry mass 

plus inert mass) resulted in 2,482 kg for the Taurus II design. 

Table 6. AXTAR mass summary for the Taurus II configuration.

AXTAR Taurus II Design MEL

Total Mass 

(kg)

1.0 Structure

2.0 Propulsion

3.0 Power

4.0 Avionics/control

5.0 Thermal control

6.0 Contingency

Dry Mass

7.0 Nonpropellant fluids

8.0 Payload/science instruments

Inert Mass

Total Less Propellant

9.0 Propellant (hydrazine)

Gross Mass

770.00

65.69

142.04

188.53

38.80

361.52

1,566.57

4.45

911.20

915.65

2,482.22

207.94

2,690.16
14

6.3 Propulsion

The propulsion system’s primary function is to deorbit the spacecraft at the end of the mission (see sec. 5.4) and to provide attitude control during each maneuver. A simple monopropellant 

blowdown system, with maximum use of off-the-shelf components, is selected for this task. The 

propulsion systems for both the Taurus II and Falcon 9 configurations are very similar, the only 

differences being the selected propellant tanks and the structural support mass. The preliminary 

system schematic is shown in figure 9. Both propulsion options consist of three diaphragm tanks 

that are loaded with hydrazine propellant and nitrogen pressurant. The thruster configuration 

includes four pods, each containing one Aerojet MR-104A/C (100 lbf) engine and two MR-106L 

(5 lbf) engines. This system provides redundancy with two of the four larger thrusters designated 

as backup. The chosen engines are a compromise between maintaining low-gravity losses during 

the deorbit burns and minimizing the engine and feed system sizes.

P

P

P

T T T

T

N2

N2H4

P

P

N2

N2H4

P

P

N2

N2H4

P

T

T

T

T P

T

T

T

T P

T

T

T

T P

T

T

T

T

Service Valve

Latch Valve

Filter

Flow Control Orifice

Temperature Sensor

Pressure Transducer

5 lbf (MR-106L)

100 lbf (MR-104A/C)

Figure 9. AXTAR propulsion system schematic.

Table 7 lists the mass breakdown for the Taurus II option propulsion system. After applying 

a conservative 30% growth allowance, the total system dry mass is 85.39 kg. An off-the-shelf tank 

is selected from Alliant Techsystems, Inc. (ATK) (model number 80488-1); the three tanks hold 

210.04 kg of hydrazine and 2.35 kg of gaseous nitrogen. The available maneuver propellant is 

207.94 kg, which gives a propellant margin of 6.7% above the required 194.80 kg (see sec. 5.4). 

The assumed tank pressure range is from 420 psia to 110 psia.
15

Table 7. Propulsion system mass summary for the Taurus II option.

Qty Item

Unit Mass 

(kg)

Total Mass 

(kg) Info

4

8

3

3

10

19

3

3

9

3

1

1

Axial thrusters

Lateral thrusters

Propellant tanks

Pressurant fill/drain valve

Pressure transducers

Temperature sensors

Propellant filters

Flow control orifice

Latch valves

Propellant fill/drain valve

Lines and fittings

Structural mounts 

1.86

0.59

11.02

0.21

0.28

0.10

0.30

0.02

0.50

0.21

4.40

4.58

7.44

4.72

33.05

0.63

2.80

1.98

0.90

0.06

4.50

0.63

4.40

4.58

Aerojet MR-104A/C

Aerojet MR-106L

ATK (80488-1)

Moog (50-856)

Lunar prospector

FCI (AS-TT)

VACCO (F1D10559-01)

AIAA 2003-4470

Moog (51-134)

Moog (50-856)

Estimate

Estimate 

Total dry mass

Total mass after 30% contingency:

65.69

85.39

6.4 Communications

At present, the downlink data rates for this mission are judged not to be in excess of fixed 

antenna capabilities. A COMM link to ground using fixed antennas is desirable over an active 

pointing design due to the elimination of gimbaled mechanisms. This results in a more reliable 

and lower risk design. Analysis shows that omnidirectional antennas are sufficient for the given 

data rates and assumed mission scenarios by using an X-band system for the science data downlink. The use of the Tracking and Data Relay Satellite System (TDRSS) for normal operations 

and data link is not desirable from a cost perspective. However, a TDRSS link during launch and 

startup operations is desirable. The COMM system designs are single fault tolerant.

For this mission, low-orbit inclinations are better for science data collection. A ground link 

analysis based on link times and daily accesses was performed to determine the best selection of 

ground stations at both 5° and 28°. South Point, Hawaii and Kourou, Guiana were selected as the 

primary and secondary ground link stations respectively for 28° orbits out of 11 possible ground 

stations analyzed. These same two stations are also selected at a 5° orbit, where Kourou becomes 

the primary station and South Point becomes the secondary station.

At these two stations, at least an 8-min primary link is possible seven times a day. Under 

a worst-case condition, a 6.7-min link five times a day is possible for the 5° inclination at the secondary South Point station. Using these two conditions as bounds, a link budget was performed 

assuming at least four links per day for 8 min each. For these link times, a 20-W transmitter can 

achieve a 90-Mbps data downlink rate. At this rate, almost 44 Gbits can be transmitted per station pass, or 172 Gbits per day. Assuming a continuous average data rate of 700 kbps for the entire 

LATA system and 72 kbps for an eight-unit SM cluster, enough link capability is left to download 

over six 15-min peak (19-Mbps) LATA events per day. Major SM events can be alternated with 

LATA event data, assuming the SM detects events first and then the LATA is pointed to the event. 

The suggested 20-W L3 X-band transmitter is presently at Technical Readiness Level (TRL) 6.
16

An estimate for the total spacecraft telemetry COMM data rate using S-band is 60-kbps 

downlink and 4-kbps uplink. Using a 5-W S-band transmitter, a link to the TDRSS for launch and 

startup operations can be accomplished at these data rates using omniantennas. The 5-W transmitter can also link with ground for normal telemetry with plenty of margin available. The suggested 

AeroAstro 5-W S-band transmitter is at TRL 8.

6.5 Avionics, Guidance, Navigation, and Control

The mass summary of the avionics system components is listed in table 8. Two fully redundant Proton 200 flight computers from SpaceMicro are the core of the avionics system. The computers are used for spacecraft operations and data management. They receive processed science data 

from the IDS and transfer the data to either the onboard data recorders or the X-band transmitters 

for downloading. The two data recorders from Surry Satellite Technology can store up to 256 Gbits 

of data at a rate of 150 Mbps. One day of required data storage is estimated to be about 173 Gbits, 

providing about 50% in memory margin. The flight computers are radiation hardened to a 100-krad 

total ionizing dose and a 70 MeV-cm2/mg single-event latch up. The Proton 200, currently at TRL 6, 

is scheduled to launch during 2011 on board the Operationally Responsive Space Satellite 1 (ORS 

Sat-1). The Surrey data recorders are at TRL 8.

Table 8. Avionics mass results for the Taurus II configuration.

Avionics System 

Components

Total Mass 

(kg) Comments

Attitude control system

Command and data systems

Instrumentation

Communications systems

Avionics cabling

95.38

14.00

15.00

30.15

34.00

Includes reaction wheels and torque rods.

Includes computers and recorders.

Includes sensors and cabling.

Includes X-band and S-band.

Power cabling not included.

Total dry mass

Total mass after 30% contigency:

188.53

245.09

Attitude knowledge is achieved using a redundant pair of Ball Aerospace star trackers and 

Northrop Grumman inertial measurement units (IMUs). The star trackers provide 4 arc-s of accuracy, meeting the 5 arc-s mission requirement. Both the IMUs and the star trackers are at TRL 8 

or above.

While this satellite has large surface areas resulting in significant disturbance torques, the 

slewing and pointing requirements are modest. Off-the-shelf reaction wheels should be sufficient 

for attitude control, keeping the cost down. In low-Earth orbit, magnetic torque rods are good 

candidates for attitude control assist and desaturation of reaction wheels. By using magnetic torquers, the spacecraft reaction control system (RCS) will not normally be required for attitude control 

and is considered for contingency purposes and disposal only. A set of three dual-coil Microcosm 

MT400-2 magnetic torquers is suggested for this mission. These torquers are at or above TRL 8.

The trade space for reaction wheels included four different wheel types at four slewing speeds. 

This trade resulted in a Teldix RSI 68-170 reaction wheel being selected. These wheels are mounted in 

a four-wheel pyramid configuration for best performance while maintaining one-fault tolerance. They 
17

have sufficient momentum and torque capability to exceed the fast slew requirement of 180° in 30 min 

while providing the desired pointing accuracy of 1 arc-min. The Teldix wheels are at or above TRL 8.

Inertial pointing will be the typical pointing mode for science observations. In this mode, the 

magnetic torque rods provide good torque authority. Using the torquers to offset the disturbances, 

the 28-hour continuous pointing goal can be achieved with unlimited pointing times possible. The 

disturbances will usually be small in a zenith-pointing mode, which is a possible scanning mode. In 

this pointing mode, desaturation of the reaction wheels can be accomplished using the magnetic 

torque rods. Atmospheric and gravity torques are at a continuous maximum in a worst-case torque 

scenario. The reaction wheels will eventually saturate if left in this mode for extended periods; 

however, extended pointing times in this mode are unlikely since there are no apparent science 

advantages. If wheel saturation does occur, a return to zenith pointing for quick desaturation is possible.

6.6 Power

PowerPower required for the AXTAR spacecraft subsystems and science instrument payload is listed in table 9.

Table 9. Power budget for the AXTAR spacecraft.

Subsystem/Element

Power 

(W)

Attitude control system

Command, data, instrumentation

Communication power

Large area timing array

SM

Instrument data system

30% power margin

152.4

125.2

179.2

600.0

121.0

40.0

365.3

Total 1,583.1

The power system was designed to meet the 5-yr mission duration assumption and operate 

in a circular orbit of at least 500-km altitude (resulting in a maximum eclipse period of 43.9 min, 

including lunar eclipse and a minimum light period of 58.8 min). Regarding the environment, a 

solar power density of 1,370 W/m2, solar panel operating temperature of 76 °C, a thermal sink 

temperature of 279 K, and a maximum ambient electronics temperature of 30 °C was assumed. 

The secondary batteries were selected based on a minimum of 25,000 charge/discharge cycles, a 

40% maximum depth of discharge, a maximum operating temperature of 30 °C, and a packing 

factor of 1.35 (i.e., battery mass is 1.35 times the total mass of all the cells). The solar arrays were 

sized assuming a 5% knockdown for cell mismatch and interconnection, a 3% per year degradation 

(due to UV radiation, thermal cycling, and contamination), and 90% cell coverage for each panel. 

The assumed bus voltage was 28 V, with a charge efficiency and discharge efficiency of 90%.

Because of the relatively long science mission duration and the high levels of sunlight available, a solar-based power system was chosen for this spacecraft. Figure 10 illustrates the basic 

elements of this design. The power system mass summary is listed in table 10.
18

Array 

Interconnect

Solar 

Array 

Panel

Solar 

Array 

Panel

SASM SASM

To Loads Switch Cards

Batteries Battery Charge 

Control

Battery Charge 

Control

Figure 10. Solar-based power system.

Table 10. Power system mass summary.

Qty Power System Components 

Unit Mass 

(kg)

Total Mass 

(kg) Comments 

2

2

2

4 

2

1

2

1

1

3

Rigid solar array wing

Solar array structure and yoke

Solar array switch module

Distribution switch 

Pyro controller

Umbilical controller

Battery charge/discharge controller 

Power system electronics enclosure

Cabling and harness 

Secondary battery

16.60

17.20

1.14

1.14 

1.10

1.14

1.15

16.70

8.00

12.42

33.20 

34.40 

2.28

4.56 

2.20

1.14

2.30

16.70

8.00

37.26

8.27 m2 area

Includes articulation mechanism. 

Orion/Mars recon orbiter

Orion

Orion

Orion

Mars recon orbiter

Scaled from COTS AITECH.

Estimate

Sized from Saft VL 48E cells.

Total dry mass

Total mass after 30% contingency

142.04

184.65

The required power generation capacity of the solar arrays is the sum of the overall power 

requirement as detailed in table 9. The overall required power amounts to 3,708 W, which consists 

of the power required to charge the secondary batteries that supply this power requirement while 

the craft is in the eclipse (dark) portion of its orbit and the losses incurred in the power system. In 

order to minimize risk, a standard, rigid panel solar array that uses spring tension to unfold during 

deployment was chosen. The total area of the arrays, including hinges and mechanisms, is 16.6 m2
19

(8.3 m2 per wing) with an areal density of 2.16 kg/m2 and a specific power (end of life) of 

241.6 W/m2. The cells are GaAs triple junction—Spectrolab ITJ or equivalent—with a 24% efficiency at 76 °C. The arrays are constructed with a ZnO front contact and a Cu back contact on 

M55J Al-honeycomb substrate.

The power electronics is comprised of a set of 11 VME circuit boards caged in a rugged 

VME enclosure that is scaled from an existing space-qualified design. Instead of specifying a particular manufacturer’s board for each function, the system is sized from representative boards using 

typical masses, power requirements, and specifications from previous missions. Each of the two 

solar array switchboards (sequential shunt) handles both array wings for full fault tolerance. Two 

pyrotechnic (pyro) event controllers manage pyro events, each handling up to eight events. For 

power distribution, each of the 4 distribution switches can control up to 16 circuits. The two battery charge controllers can charge two secondary batteries each, and the single umbilical switch can 

switch up to eight external high-current circuits. The VME enclosure is sized with fully redundant 

power supplies.

Because the spacecraft and science instrumentation must be powered during those periods 

of the orbit in which the solar arrays get no sunlight, energy must be stored during the lighted 

periods. Li-ion secondary batteries were chosen for this purpose because of their favorable energy 

density and high reliability. The battery cells are Saft Li-ion VL48E cells (3.6 V, 48 Ah). Each battery contains cell-balancing electronics and consists of eight cells for a total charge capacity of 28.8 

V. The total number of battery units is three, each having a mass of 12.42 kg. The maximum depth 

of discharge is 40%.

6.7 Structures

Structural requirements of the AXTAR spacecraft bus are driven by four main factors: 

(1) Launch platform limitations; (2) primary science instruments’ shape, size, and mass; (3) the 

requirement to mount 20 LATA super modules in a coplanar array; and (4) radiation shielding. 

The Taurus II 3.45-m diameter launch platform limits LATA array width to four super modules, 

necessitating a 4×5 LATA array. Radiation shielding against cosmic and solar radiation is incorporated into the spacecraft bus. Since the LATA array is quite massive and requires heavy metal 

cosmic radiation shielding on three sides, the Taurus II CG height limit of 2 m from the payload 

interface plane limits the height of the spacecraft bus. The LATA also requires solar radiation 

shielding for viewing angles up to 45° from the Sun, which is installed in parallel planes perpendicular to the five rows of the LATA array. The resulting AXTAR-Taurus spacecraft configuration 

meets the Taurus II payload static envelope, CG, and payload mass-to-orbit estimate for the mission profile, as well as the science instruments’ performance requirements with positive structural 

margin of safety. See figure 8 for configuration features and details.

The spacecraft concept uses lightweight 2024–T351 Al panels, tubing struts, and T- and 

I-beam structural supports for component and science instrument mounting. The panels conduct 

heat and double as radiators for thermal management. Midway fore and aft on opposite sides of 

the spacecraft bus, two flight-proven telescoping booms support the fold-out solar arrays and are 

stowed against the spacecraft bus for launch. These were sized similarly to those used on the 

Hubble Space Telescope but proportionally less massive.3
20

A short, octagonal main spacecraft bus provides the needed structure and surfaces to secure 

propulsion, avionics, and other spacecraft components on two interior panels of a four parallelpanel structure. The top and bottom exterior panels provide mounting for antennas and other 

sensors, as well as 13 of the 27 SMs. The aft bulkhead panels are used for thermal management and 

support the Taurus II launch adapter ring. The four parallel panels and two perpendicular walls 

that support and separate them are one-eighth- to one-quarter-in (0.32–0.64-cm) thick to support 

all the mounted equipment and enhance thermal management. Most exterior surfaces are much 

thinner and are primarily thermal closeouts, rather than structural.

The science bus structure needed to support the LATA array and 14 SMs is forward of the 

main spacecraft bus. Graded-Z galactic radiation shielding utilizing tin/tantalum/Al is modeled 

as nonstructural mass in the science bus structure, conceptually similar to the shielding flown on 

the Swift Burst Alert Telescope (BAT).4 The availability of tin (foil) may not support the fabrication of this shielding. Copper replaced tin for the Swift BAT fabrication. This change should be 

negligible to structural mass estimates in this study. A lightweight solar shade truss structure of 

0.04–0.1 gauge (0.09–0.25-cm thick) round tubing struts bears six lead/Al Sun shades, tall enough 

to shade each super module.

A Finite Element Modeling Analysis and Postprocessing (FEMAP) model was created 

based on the Pro/E configuration, loaded with components and instrument masses, and analyzed 

with NX NASTRAN. Eighteen load cases were run with a load set of combined Taurus II/Falcon 9 launch vehicle loads, +6/–2 g (acceleration due to gravity at standard sea level) axial and ±2 g 

lateral, at 30° and 45° intervals. A 1.4-safety factor for isotopic strength and a 0.65-buckling factor were used in the analysis. Once the structural analysis was successful, the model was optimized 

using Hypersizer® for minimum mass. Structural members were grouped, results were analyzed, 

and material thicknesses were adjusted to reflect common raw material stock and other manufacturability criteria. The results were iterated through FEMAP until a minimum mass, positive 

margin of safety (MOS), and total deformation of less than 3 cm was converged upon using a compressive yield stress of 40,000 psi (2.76×108 Pa) for Al. The FEMAP model used to analyze and 

size the spacecraft is depicted in figure 11. Plates, bars, and nonstructural and rigid-mass 

elements are used to represent all masses estimated by all the disciplines developing the spacecraft.

Table 11 summarizes the structural mass of the spacecraft bus, science bus, shielding,and 

secondary structure. The AXTAR-Taurus concept structure has a high MOS (i.e., low stresses) 

because some panels were thickened to limit excessive bending displacements and maximize 

thermal conduction and radiation to space. Total mass after 30% contingency is 1,001 kg.
21

Bending Displacements <3 cm

Sun Shades (6)

Figure 11. FEMAP Taurus II model deformed side view showing rows 

of Sun shades and exaggerated deformations.

Table 11. AXTAR structures mass summary for the Taurus II configuration.

Structural Components 

Mass 

(kg) Comments 

Spacecraft bus 

Science/LATA bus 

Galactic radiation shielding 

Solar shade supports 

Solar radiation shielding 

Secondary structure 

215.5 

164.0 

306.0 

30.5 

25.0 

29.0 

Aft end housing spacecraft systems and launch adapter. 

Forward end of spacecraft supporting 20 LATAs. 

Nonstructural mass shielding three sides of all LATA super modules. 

Nonstructural components supporting solar shielding. 

Nonstructural mass shielding all LATA super modules in the spacecraft X-Z plane. 

SM mounting structures.

Total dry mass 

Total mass after 30% contingency

770.0 

1,001.0

6.8 Thermal

A passive-thermal design concept was developed for the AXTAR spacecraft, passive meaning that no actively pumped fluid systems are required. Thermal control of the AXTAR spacecraft 

will use components including multilayer insulation, high emissivity paint and coatings, heaters, 

etc. to maintain spacecraft subsystem components within acceptable temperature ranges. There are 

no dedicated radiators; spacecraft structural panels act to dissipate avionics heat by conduction 

and also act as radiative surfaces. The bus outer surfaces are covered in low absorptivity materials 

in order to cold bias the spacecraft and minimize temperature fluctuations due to orbital position. 

Propellant tanks are wrapped in multilayer insulation (MLI) and propellant tank heaters are sized 

for worst-case cold orientation. RCS thrusters, antennas, solar arrays, and solar array mechanisms 

are not part of the prephase A analysis.
22

A system-level thermal model of the spacecraft bus and LATA structure was developed 

using Cullimore and Ring Technologies’ Thermal Desktop® (<http://www.crtech.com>) to assess 

spacecraft bus and LATA support structure interface temperatures. The model geometry is shown 

in figure 12 and represents the Taurus II configuration. The structure is modeled as Al, the panel 

thickness being consistent with the structural design. Environmental heat loads were calculated for 

an Earth orbit altitude of 585 km. Steady-state spacecraft structure temperatures for both a hot 

and cold orientation were generated in order to determine a range of temperatures to be expected 

during the on-orbit operations. Full subsystem and experiment heat loads were applied. The hot 

orientation is defined as LATA to a Sun angle of 30° with a beta angle of 50°, and the cold 

orientation is defined as LATA to a Sun angle of 90° with a beta angle of 0°. Both hot and cold 

orientations are shown in figure 13.

Sun Shades (6)

Figure 12. Thermal model geometry for the Taurus II configuration.

Cold Case, Beta = 0o Hot Case, Beta = 50o

Sun Angle = 30o

(a) (b)

Figure 13. Orientation of spacecraft showing the (a) cold case and (b) hot case 

used in the thermal analysis.

The spacecraft bus internal surfaces are assumed to be black anodized to optimize radiative 

exchange within the enclosure. White paint is used on the exterior surfaces to provide a low absorptivity (α)to emissitivity (ε) ratio (α/ε)that serves to minimize structure temperatures. The backside 

of the LATA support structure is covered with a 12-layer MLI blanket, with the outer layer being 

beta cloth painted white. LATA supports and Sun shades are also painted white. Optical properties 

used for the thermal model surfaces are itemized in table 12 and were taken from various sources 
23

including the Spacecraft Thermal Control Handbook5 and NASA Spacecraft Thermal Coatings 

Reference.6 MLI ε* values represent effective emissivity for the blanket.

Table 12. Thermal model surface optical properties.

Material Absorptivity Emissivity

Spacecraft bus internal surfaces

Spacecraft bus external surfaces

Spacecraft bus closeouts

LATA support structure

LATA supports

LATA sun shades

RCS tanks

Black anodized

White paint

White paint on beta cloth

Inner layer, black kapton® (MLI=5 layers) ε*= 0.02

White paint on beta cloth (12 layers) ε*= 0.004

White paint

White paint

MLI (5 Layers, AlK), ε*= 0.02

0.90

0.25

0.17

0.92

0.17

0.25

0.25

0.60

0.90

0.87

0.92

0.88

0.92

0.87

0.87

0.09

Subsystems equipment and experiment heat loads assumed for the thermal analysis are 

shown in table 13 for the Taurus II configuration. A total of 1,566 W of power/heat dissipation 

was considered in the thermal analysis. All heat loads are imposed directly on the structure and 

modeled as area averaged heat loads. Heat loads were distributed on the spacecraft panels according to specific box locations. Temperature predictions reflect structure interface temperatures and 

are documented in table 13. Experiment temperature prediction is not included as part of the subsystem thermal analysis. Each of the three propellant tanks will require a maximum of 3-W heater 

power, as sized by the cold orientation. All predicted interface temperatures are within the 

acceptable range, as shown in table 13.

Table 13. Heat dissipation, operating, and predicted temperatures 

for the Taurus II configuration.

Total Heat 

Dissipation 

(W)

Operating 

Temperature Range 

(°C)

Predicted Interface 

Temperature Range 

(°C)

ACS control system

Reaction wheels (4)

Magnetic torquers (3)

Command and data system

COMM system

Power systems enclosure

Batteries (3)

RCS tank heaters (3)

IDS

LATA (20)

SM (27) + telemetry hub 

Total

43.4

97.6

11.4

125.2

179.1 

167.0 

185.4 

9.0

40.0 

600.0 

117.0

1,566.1* 

–40 to 85

–20 to 60

–43 to 66

–40 to 85

–20 to 70

–40 to 85

0 to 45

10 to 30

–40 to 5

–40 to 10

–40 to 10

–

–10 to 15 

–10 to –7 

–10 to –3 

–10 to 15 

–10 to 15 

–10 to 23 

7 to 23 

19 to 25 

–20 to –11 

–8 to 0 

–30 to 3 

–

* This total does not include RCS tank heaters since this is an intermittent requirement.

Steady state analysis results for the spacecraft bus and LATA support structure cold case 

are shown in figures 14 and 15, respectively. Structure temperatures range from 25 °C to 35 °C 

for the bus and –8 °C to –67 °C for the LATA structure. The steady state analysis results for the 
24

spacecraft bus and LATA support structure hot case are shown in figures 16 and 17, respectively. 

Structure temperatures range from –15 °C to 23 °C for the bus and 4 °C to –32 °C for the LATA 

structure. The spacecraft bus and LATA support structure were analyzed as a single unit, although 

temperature results are shown separately in figures 14 through 17 for clarity. Environmental heat 

loads are averaged about an orbital period. All predicted interface temperatures are within 

acceptable operating ranges for as-defined hot and cold orientations. 

>25.2

25.2

19.2

13.3

7.28

1.3

–4.68

–10.7

–16.6

–22.6

–28.6

–34.6

Node

7 SM and Hub Power Systems and Batteries

Rx Wheels (4)

RCS Tanks

6 SM and IDS

COMM, C&DS, ACS

Steady State, Beta 0, Orbital Averaged Heat Rates 

(Temperatures in °C), 100% Heat Dissipation

Figure 14. Analysis results for the Taurus II configuration: Spacecraft bus (cold case).

>–7.85

–7.85

–13.8

–19.7

–25.7

–31.6

–37.6

–43.5

–49.5

–55.4

–61.4

–67.3

Node

4 SM

4 SM

6 SM

Sun Shades (6)

Steady State, Beta 0, Orbital Averaged Heat Rates 

(Temperatures in °C), 100% Heat Dissipation

Figure 15. Analysis results for the Taurus II configuration: LATA structure (cold case).
25

>22.9

22.9

19.1

15.3

11.5

7.7

3.89

0.0738

–3.74

–7.55

–11.4

–15.2

Node

Steady State, Beta 50, Orbital Averaged Heat Rates 

(Temperatures in °C), 100% Heat Dissipation

RCS Tanks

Figure 16. Analysis results for the Taurus II configuration: Spacecraft bus (hot case).

Sun Shades (6) >3.83

3.83

0.201

–3.43

–7.05

–10.7

–14.3

–17.9

–21.6

–25.2

–28.8

–32.4

Steady State, Beta 0, Orbital Averaged Heat Rates 

(Temperatures in °C), 100% Heat Dissipation

Node

Figure 17. Analysis results for the Taurus II configuration: LATA structure (hot case).

Thermal management is accomplished with typical, flight proven components, so no 

technology development is required. Total thermal control mass for the Taurus II configuration 

is estimated at 50.4 kg, which includes a 30% margin as shown in table 14.
26

Table 14. Thermal subsystem mass summary for the Taurus II configuration.

Thermal System Components

Total Mass 

(kg) Comments

Multilayer insulation/thermal tape

Thermal filler

Paint/thermal coatings

Heaters/thermostats

31.0

1.2

6.1

0.5

10–12 Layer blanket

Chotherm 

White paint

–

Total dry mass

Total mass after 30% contingency

38.8

50.4
27

7. PRELIMINARY SPACECRAFT DESIGN FOR THE FALCON 9 LAUNCH VEHICLE

The science team was interested in knowing the maximum number of LATA super modules 

that could be accommodated on a spacecraft designed to use the higher performance of a larger 

launch vehicle such as the Falcon 9. Therefore, the spacecraft design team created an alternative 

configuration for that vehicle. While time constraints during the study prevented the team from 

iterating until the design closed, the mass budget and subsystem designs are representative of what 

would have been the final design had the study continued.

7.1 Configuration

The driving factor in the spacecraft’s configuration was the number of the primary science 

instruments (LATA super modules) desired. The Falcon 9 shroud size would allow for more super 

modules to be carried than were needed to accomplish the science goals. The LATA’s coplanar and 

pointing requirements drove the design layout to be divided into two main sections, a spacecraft 

system bus and a science bus (see fig. 18). The science bus/LATA array was set forward to provide 

viewing separation from the spacecraft’s solar arrays and other systems. The science bus structure 

was designed to allow the LATA super modules to be nestled inside the primary structure to provide additional shielding. With a relatively large shroud diameter, it was possible to place six super 

modules in the spacecraft bus section for a more integrated compact vehicle. To help meet the Sun 

avoidance-angle plane requirements, a separate Sun-shade secondary structure was placed above 

the LATA in addition to the primary vehicle structure. This gave the required shielding height 

necessary for the rows of LATA super modules.

SM Cluster (7)

Solar Array (2×)

42 LATAs

Sun Shields

SM Cluster (6)

Figure 18. Configuration of the AXTAR spacecraft: Falcon 9 option.
28

Another design factor was the number and placement of the SM cameras. A 32-face 

dodecahedron would allow full-sky coverage. These were broken up into several clusters and placed 

on the spacecraft to optimize view angles while minimizing size. These camera clusters were fairly 

low mass, so no additional provisions were needed to accommodate them. Final placement and 

grouping would need to be optimized in a further study.

The spacecraft system bus is located aft near the launch vehicle interface to carry launch 

vehicle loads and to allow for the spacecraft’s systems to be enclosed in a more compact volume. 

The central core of the spacecraft bus houses the propulsion tanks and the control moment gyros 

(CMGs). Discussion of the use of CMGs is in section 7.4. The outer bays/walls allow for mounting of the other avionics and spacecraft components. The design has continuous load paths and 

structures, allowing for minimum mass to be obtained. Locating the spacecraft bus aft also allows 

for desirable CG characteristics. The overall configuration is conservative and does allow room for 

component growth and for extra subsystem components to be added that were not analyzed in this 

study. The design based on the Falcon 9 launch vehicle could be used on comparable (not smaller) 

size launch vehicles.

7.2 Mass Properties

The mass breakdown for the overall vehicle, spacecraft subsystems, and science instruments 

for the Falcon 9 design is presented in table 15. The Falcon 9 design resulted in a total vehicle gross 

mass of 4,972 kg. Gross mass is the combined total of vehicle dry mass, inert mass, and propellant. Dry mass is defined as spacecraft subsystems mass minus the useable propellant, propellant 

residuals (see item 7.0 in table 15), and science instruments. The dry mass total, including the 30% 

contingency, resulted in a mass of 2,766 kg for the Falcon 9 design. Inert mass includes propellant 

residuals and science instruments. The inert mass for the Falcon 9 design totaled 1,801 kg. The 

total less propellant (dry mass plus inert mass) resulted in 4,567 kg for the Falcon 9 design. 

Table 15. AXTAR mass summary for the Falcon 9 configuration.

AXTAR Falcon 9 Design MEL

Total Mass 

(kg)

1.0 Structure

2.0 Propulsion

3.0 Power

4.0 Avionics/control

5.0 Thermal control

6.0 Contingency

Dry Mass

7.0 Nonpropellant fluids

8.0 Payload/science instruments

Inert Mass

Total Less Propellant

9.0 Propellant (hydrazine)

Gross Mass

1,330.00

94.66

226.31

422.53

53.90

638.22

2,765.61

4.09

1,797.20

1,801.29

4,566.90

405.25

4,972.15
29

7.3 Propulsion

The mass breakdown for the Falcon 9 option propulsion system is listed in table 16. This 

system is identical to the one for the Taurus II configuration except for the propellant tanks and 

support structure (see sec. 6.3). The total predicted dry mass for this option is 123.05 kg. An ATK 

tank (model 80325-1) is also chosen for this configuration. The totals for loaded propellant and 

pressurant are 409.34 kg and 4.82 kg, respectively. The available maneuver propellant is 405.25 kg 

(11.5% margin). The assumed tank pressure range is from 420 psia to 110 psia.

Table 16. Propulsion system mass summary for the Falcon 9 option.

Qty Item

Unit Mass 

(kg)

Total Mass 

(kg) Info

4

8

3

3

10

19

3

3

9

3

1

1

Axial thrusters

Lateral thrusters

Propellant tanks

Pressurant fill/drain valve

Pressure transducers

Temperature sensors

Propellant filters

Flow-control orifice

Latch valves

Propellant fill/drain valve

Lines and fittings

Structural mounts 

1.86

0.59

20.00

0.21

0.28

0.10

0.30

0.02

0.50

0.21

4.40

6.60

7.44

4.72

60.00

0.63

2.80

1.98

0.90

0.06

4.50

0.63

4.40

6.60

Aerojet MR-104A/C

Aerojet MR-106L

ATK (80325-1)

Moog (50-856)

Lunar Prospector

FCI (AS-TT)

VACCO (F1D10559-01)

AIAA 2003-4470

Moog (51-134)

Moog (50-856)

Estimate

Estimate 

Total dry mass

Total mass after 30% contingency:

94.66

123.05

7.4 Communications, Avionics, Guidance, Navigation, and Control

For the larger Falcon-9 design, slewing is much more challenging, and CMGs are considered. Using CMGs in place of reaction wheels to accommodate the higher moments of inertia and 

disturbance torques provides the required performance with margin. Although the mass of CMGs 

is moderately higher than reaction wheels, the power demand is greatly reduced.

Using a pointing dish antenna to accommodate the higher data rates for an additional 22 

LATA modules (basically doubling the data rate) greatly reduces the power requirement for transmission. However, this adds a pointing mechanism to the design. This pointing mechanism is estimated to be at TRL 8. For the additional memory requirement and torque authority needed, the 

number of data recorders and magnetic torquers were doubled. This resulted in a relatively small 

increase in mass and power. The mass summary is listed in table 17.
30

Table 17. COMM, avionics, and GN&C mass summary 

for the Falcon 9 configuration.

Avionics System 

Components

Total Mass 

(kg) Comments

Attitude control system

Command and data systems

Instrumentation

Communications systems

Avionics cabling

308.98

20.00

15.00

38.55

40.00

Includes reaction wheels and torque rods

Includes computers and recorders

Includes sensors and cabling

Includes X-band and S-band

Power cabling not included

Total dry mass

Total mass after 30% contingency

422.53

549.29

7.5 Power

The extended configuration for the Falcon 9 had power requirements as shown in table 18.

Table 18. Power budget for AXTAR for the Falcon 9 configuration.

Subsystem/Element

Power 

(W)

Attitude control system

Command, data, instrumentation

Communication power

Large area timing array

SM

Instrument data system

30% power margin

174.2

178.8

120.3

1,260.0

129.0

40.0

570.7

Total 2,473.0

Using the same power system design, the team resized the components for the new power 

requirement. Table 19 summarizes the result.

Table 19. Power system mass summary for the Falcon 9 configuration.

Qty Power System Components 

Unit Mass 

(kg)

Total Mass 

(kg) Comments 

2

2

3

4

2

1

3

1

1

6

Rigid solar array wing

Solar array structure and yoke

Solar array switch module

Distribution switch

Pyro controller

Umbilical controller

Battery charge/discharge ctl 

Power system electronics enclosure

Cabling and harness 

Secondary battery

24.96

25.89

1.14

1.14

1.10

1.14

1.15

18.32

17.00

12.42

49.92

51.78

3.42

4.56

2.20

1.14

3.45

18.32

17.00

74.52

11.6 m2 area

Includes articulation mechanism 

Orion/Mars recon orbiter

Orion

Orion

Orion

Mars recon orbiter

Scaled from COTS AITECH

Estimate

Sized from SAFT VL 48E cells

Total dry mass

Total mass after 30% contingency

226.31

294.20
31

7.6 Structures

The Falcon 9 larger, 4.6-m diameter launch platform can launch a much larger spacecraft 

with greater science capability. Over twice as many LATA super modules were accommodated 

while maintaining a low CG. Spacecraft structural mass was scaled from the AXTAR-Taurus 

configuration (which also meets all Falcon 9 launch platform requirements) to accommodate the 

greater mass of 42 LATA super modules and other spacecraft components. An estimated 70% 

increase in structural mass is necessary to accommodate the larger and more capable AXTARFalcon concept. The FEMAP model used to analyze and size the spacecraft is depicted in figure 19.

(a) (b)

SMs (7)

SMs (6)

Sun Shields (9)

Figure 19. FEMAP model used for the Falcon 9: (a) Side view and (b) top view.

Table 20 summarizes the structural mass of the AXTAR-Falcon concept structure. Total 

mass after 30% contingency is 1,729 kg.

Table 20. Structural mass summary for the AXTAR spacecraft 

for the Falcon 9 configuration.

Structural Components 

Mass 

(kg) Comments 

Spacecraft bus 

Science/LATA bus 

Galactic radiation shielding 

Solar shade supports 

Solar radiation shielding 

Secondary structure 

323 

246 

612 

65 

55 

29 

Aft end housing spacecraft systems, launch adapter, and 12 LATAs 

Forward end of spacecraft supporting 30 LATAs 

Nonstructural mass shielding three sides of all LATA super modules 

Nonstructural components supporting solar shielding 

Nonstructural mass shielding all LATA super modules in the spacecraft X-Z plane 

SM mounting structures

Total dry mass 

Total mass after 30% contingency

1,330 

1,729
32

7.7 Thermal

Thermal control for the Falcon 9 configuration is consistent with the smaller Taurus II 

design. Mass estimates were determined by appropriate scaling to account for a larger bus and support structure and are presented in table 21. The total mass after a 30% contingency is estimated 

at 70.1 kg. Table 22 details the heat dissipation and operating temperatures for the subsystems 

and experiment components. Although no analyses were performed for this configuration, it is 

expected that all structural temperatures will be acceptable. Battery location adjustments may be 

required to accommodate the increased heat dissipation. A total of 2,362 W of spacecraft power/

heat dissipation is estimated for this conf﻿iguration. 

Table 21. Thermal system mass summary for the AXTAR spacecraft 

for the Falcon 9 configuration.

Thermal System Components

Total Mass 

(kg) Comments

Multilayer insulation/thermal tape

Thermal filler

Paint/thermal coatings

Heaters/thermostats

42.0

2.1

9.1

0.7

10–12 Layer blanket

Chotherm 

White paint

–

Total dry mass

Total mass after 30% contingency

53.9

70.1

Table 22. AXTAR heat dissipation and operating temperatures for the Falcon 9.

Total Heat Dissipation 

(W)

Operating Temp Range 

(°C)

ACS control system

Reaction wheels (4)

Magnetic torquers (6)

Command and data system

COMM system

Power systems enclosure

Batteries (6)

RCS tank heaters (3)

IDS interface to bus

LATA (42) interface to structure

SM (28) + telemetry hub interface

43.4 

108.0 

22.8 

178.8 

120.3 

189.0 

279.0 

12.0

40.0 

1,260.0 

121.0 

–40 to 85 

–20 to 60 

–43 to 66 

–40 to 85 

–20 to 70 

–40 to 85 

0 to 45 

10 to 30 

–40 to 5 

–40 to 10 

–40 to 10 

Total 2,362.0* –

* This total does not include the RCS tank heaters since this is an intermittent requirement.
33

8. RISK ANALYSIS

In identifying risks for both the spacecraft and the science instrumentation, a taxonomic 

approach was chosen. The risks were identified from an analysis of the uncertainties discovered 

in the following areas:

• Requirements: How certain and well defined are the science objectives and the requirements 

derived from them?

• Design: How certain are the design parameters and design decision consequences?

• Testing: How defined and complete are the plans for testing?

• Programmatic considerations: How certain are the funding program’s guidelines, limitations, 

and mission requirements and the team’s ability to satisfy them?

The individual risks were identified through interviews with each of the study participants. 

Each risk was then evaluated in terms of its likelihood (as a probability), its impact (in dollars), 

and possible options for management (mitigation, acceptance, transfer, etc.) at the next study level. 

There were a number of risks identified and they may be summarized as follows:

• Mission requirements are not well defined. No formal requirements analysis has been 

performed. Performing a formal requirements analysis will fully mitigate those risks.

• Electronics associated with the x-ray detectors have not been designed and power consumption is far lower than any previously implemented functional equivalent. There is no reason to 

believe that the power consumption requirement can be met; however, a risk impact analysis 

performed to determine the impact of higher power consumption on the spacecraft showed that 

the impact is minimal (a 50% higher consumption will result in about 30 kg of power system 

mass). Therefore, relaxing the requirement is suggested.

• Currently, there are no test plans for any of the science instrumentation, and no known similar 

test plans are available for adaptation. Consequently, the facilities available for testing along 

with the cost and schedule for testing is unknown. A formal test planning exercise will fully 

mitigate this risk.

• There are materials availability issues with high-purity Si and tin. These may be mitigated 

programmatically by advanced planning.
34

• Ultimate costs of both target launch vehicles are unknown, and there may not be an adequate 

vehicle available at launch date in the Medium-Class Explorer (MIDEX) cost range. This risk 

is common to most MIDEX missions and must be managed as testing of these vehicles progresses. A lighter version of AXTAR capable of fitting on a Minotaur IV would provide 

a mitigating contingency.
35

9. ANIMATION

The purpose of creating the AXTAR animation was to introduce the spacecraft design 

and its components to space science viewers. The animation focused primarily on the spacecraft 

instruments and LATA super module functionality. Textual labeling was used to highlight the 

instruments with technical information pertinent to their target. Additionally, the labeling served 

as basic narration for viewers.

The LATA super module sequence was depicted in greater detail than other instruments. 

An exploded view of this science instrument showed x rays represented by glowing particles being 

selected by the collimator and striking the Si pixel detector. The underlying interposer board was 

shown creating a charge. The sequence continued as the interposer board charges were converted 

into digital data and sent to the computer. This animated sequence is a comprehensive representation of LATA component assembly and functionality.

The basic processes used for this animation were modeling, texturing, lighting, animating, 

rendering, and compositing. The compositing process was somewhat robust for this animation. 

Each instrument and component was rendered on individual layers. While this process is more time 

consuming, it provides animators with more control of final appearances. Multiple layers of instruments, backgrounds, and text were composited together using Autodesk Combustion 2008. The 

entire AXTAR animation contains over 30 layers of elements.

The use of professional animation software and processes resulted in a complete AXTAR 

animation. The video can be shown at conferences and used in presentations.
36

10. CONCLUSIONS

The spacecraft design team completed two spacecraft concepts for the AXTAR mission—

one with a minimal set of science instruments and one with a much larger set. While the design of 

the smaller configuration was driven to closure, the team had the time and resources to complete 

just one design iteration of the larger configuration. Since the main focus of the study was to 

complete the design of the smaller configuration and the larger configuration was a secondary 

objective, all study goals were met.

A summary showing the major elements of each spacecraft configuration is presented in 

table 23. The smaller configuration, which contains 20 LATA super modules and 27 SM cameras, can be launched on either the Taurus II (and Taurus II enhanced) or the Falcon 9 or similar 

vehicle. Launching on the Falcon 9 provides a lower final inclination, helping to minimize the SAA 

effect. The larger configuration, with 42 LATA super modules, is much more massive and requires 

a Falcon 9 or similar vehicle. For either configuration, reducing the mass will result in a lower final 

inclination. Also of note is that the spacecraft subsystems do not require the development of any 

new technologies; all subsystem components are either existing or are based on existing technology.

Table 23. Summary data for the two spacecraft designs.

Smaller (Taurus II) 

Configuration

Larger (Falcon 9) 

Configuration*

LATA super modules (qty)

SMs (qty)

Total observatory mass (kg)

Final orbit inclination with Taurus II /Taurus II 

 Enhanced, launched from CCAFS (deg)

Approximate final orbit inclination with Falcon 9 

 launched from CCAFS/Kwajalein (deg)

Spacecraft technologies needing development

20

27

2,700

25/18**

12/5

None

42

27–32

5,000 (est)

na/na

19/5

None

* These are preliminary results; design is not closed.

** Taurus II enhanced performance is very preliminary.

Additional design work could uncover potential mass savings in the design, as both spacecraft designs fell on the ‘heavy’ curve of the relationship between payload mass and dry bus mass 

(see fig. 3). The primary benefit of lower mass would be a lower final orbital inclination and result 

in reducing the SAA effect. If somehow the number of LATA super modules could be reduced and 

still satisfy the science mission goals, the spacecraft could possibly be launched on a smaller vehicle, 

such as the Minotaur IV.
37

REFERENCES

1. Ray, P.S.; Chakrabarty, D.; Wilson-Hodge,C.A.; et al.: “AXTAR: Mission Design Concept,” 

Proc. SPIE, Vol. 7732, p. 773,248, 2010.

2. Hope, S.; and Vaughn, F.J.: “Trajectory Design and Control of the Compton Gamma Ray 

Observatory Reentry,” AAS 01-075, 24th Annual AAS Guidance and Control Conference, 

Breckenridge, Colorado, January 31–February 4, 2001.

3. Maly, J.R.; Pendleton, S.; Salmanoff, J.; et al.: “Hubble Space Telescope Solar Array 

Damper,” Smart Struct. Mater., CSA Engineering Inc., <maly@csaengineering.com>, 

Newport Beach, CA, 1999.

4. Robinson, D.W.: “Building a New Kind of Graded-Z Shield for Swift’s Burst Alert Telescope,” Doc. ID # 20030032274, NASA Goddard Space Flight Center, Greenbelt, MD, 

<http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20030032274_2003026279.pdf>, 

accessed October 2010.

5. Spacecraft Thermal Control Handbook, D.G. Gilmore (ed.), The Aerospace Corporation, 

The Aerospace Press, El Segundo, CA, pp. 4–79, 1994.

6. L. Kauder: “Spacecraft Thermal coating References,” NASA/TP—2005–212792, Goddard 

Space Flight Center, Greenbelt, MD, pp. 40–48, 2005.
38

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Advanced X-Ray Timing Array Mission:

Conceptual Spacecraft Design Study

R.C. Hopkins, L. Johnson, H.D. Thomas, C.A. Wilson-Hodge, 

M. Baysinger*, C.D. Maples*, P.D. Capizzo**, L.L. Fabisinski***, 

L. Hornsby†, K.S. Thompson†, and J.H. Miernik‡

George C. Marshall Space Flight Center

Huntsville, AL 35812

National Aeronautics and Space Administration

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Unclassified-Unlimited

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Prepared by the Advanced Concepts Office at Marshall Space Flight Center

*Qualis Corporation, **Raytheon Co., ***International Space Systems, Inc., †Jacobs ESTS Group, 

and ‡ERC, Inc./Jacobs ESTS Group

M–1325

Technical Memorandum

NASA/TM—2011–216476

black hole, neutron star, x-ray timing, spacecraft design

01–11–2011

UU 54

NASA

U U U

The Advanced X-Ray Timing Array (AXTAR) is a mission concept for submillisecond timing of bright 

galactic x-ray sources. The two science instruments are the Large Area Timing Array (LATA) (a collimated 

instrument with 2–50-keV coverage and over 3 square meters of effective area) and a Sky Monitor (SM), 

which acts as a trigger for pointed observations of x-ray transients. The spacecraft conceptual design team 

developed two spacecraft concepts that will enable the AXTAR mission: A minimal configuration to be 

launched on a Taurus II and a larger configuration to be launched on a Falcon 9 or similar vehicle.

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NASA’s mission.

Specialized services that complement the STI 

Program Office’s diverse offerings include creating 

custom thesauri, building customized databases, 

organizing and publishing research results...even 

providing videos.

For more information about the NASA STI Program 

Office, see the following:

• Access the NASA STI program home page at 

<http://www.sti.nasa.gov>

• E-mail your question via the Internet to 

<help@sti.nasa.gov>

• Fax your question to the NASA STI Help Desk 

at 443 –757–5803

• Phone the NASA STI Help Desk at 

443 –757–5802

• Write to:

NASA STI Help Desk

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076–1320
National Aeronautics and

Space Administration

IS20

George C. Marshall Space Flight Center

Huntsville, Alabama 35812