**DEBRIS ASSESSMENT REPORT**

**FOR THE**

**IMAGE MISSION**

November 1997

SwRI Project 15-8089

Document No. 8089-DAR-01

Revision 1

Contract NAS5-96020

Prepared by

![](media/image1.wmf){width="1.445138888888889in"
height="1.0576388888888888in"}

SOUTHWEST RESEARCH INSTITUTE

Instrumentation and Space Research Division

6220 Culebra Road, San Antonio, Texas 78228-0510

\(210\) 684-5111 • FAX (210) 647-4325

**DEBRIS ASSESSMENT REPORT**

**FOR THE**

**IMAGE MISSION**

SwRI Project 15-8089

Document No. 8089-DAR-01

Revision 1

Contract NAS5-96020

Prepared by: Mark Tapley Date: 6 November 1997

> Systems Engineer

Approved by:
\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_ Date:
\_\_\_\_\_\_\_\_\_\_\_\_

W. C. Gibson, Project Manager

Approved by:
\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_ Date:
\_\_\_\_\_\_\_\_\_\_\_\_

> J. L. Burch, Principal Investigator

Approved by:
\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_ Date:
\_\_\_\_\_\_\_\_\_\_\_\_

P. Brian Gupta, SwRI Quality Assurance

Approved by:
\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_\_ Date:
\_\_\_\_\_\_\_\_\_\_\_\_

> Frank Volpe, Mission Manager

Instrumentation and Space Research Division

Southwest Research Institute

P. O. Drawer 28510

6220 Culebra Road

San Antonio, Texas 78228-0510

\(210\) 684-5111

**TABLE OF CONTENTS**

Page

1\. BRIEF BACKGROUND ON PROGRAM AND PROGRAM MANAGEMENT

1.1 Mission Description

1.2 Program/Project Objectives

1.3 Program/Project Schedule

1.4 Responsible program or project manager

2\. DESCRIPTION OF DESIGN AND OPERATIONS FACTORS

2.1 Hardware

2.1.1 Physical description of main structure

2.1.2 Description of surfaces/materials exposed to space

2.1.3 Description of spacecraft components most sensitive to debris
impact

2.1.4 Description and location of pressurized volumes

2.1.5 Description of on-board propellants

2.1.6 Description and location of fuel storage and transport systems

2.1.7 Description of range safety systems

2.1.8 Description of systems containing stored kinetic energy

2.2 Mission Parameters

2.2.1 Number of Spacecraft

2.2.2 Launch Date and Time

2.2.3 Mission Orbit

2.2.4 Flight Attitude

3\. ASSESSMENT OF DEBRIS RELEASED DURING NORMAL OPERATIONS

3.1 Debris Released During Staging, Payload Separation, and Payload
Deployment

3.1.1 Description of debris released, size, mass, area, initial orbit,
and lifetime

3.1.2 Calculated area-time for debris greater than 1 mm (mm^2^-year)

3.1.3 Calculated object-time for debris greater than 1 mm (mm^2^-year)

3.1.4 Calculated time for removal of debris from GEO altitude to at
least 300 km below GEO altitude

3.1.5 Source for analysis

3.2 Debris Released During Mission Operations (N/A)

4\. ASSESSMENT OF ORBITAL DEBRIS GENERATED BY EXPLOSIONS AND INTENTIONAL
BREAKUPS

4.1 Explosions from On-Board Stored Energy

4.1.1 Description of failure modes leading to explosion

4.1.2 Description of systems involved in explosive failure

4.1.3 Estimated probability of explosion

4.1.4 Detailed plan for safing structure after completion of mission

4.2 Intentional Breakups (N/A)

5\. ASSESSMENT OF DEBRIS GENERATED BY ON-ORBIT COLLISIONS

5.1 Assessment of Collisions with Large Objects During Mission
Operations

5.1.1 Estimated probability of collision with intact space systems or
large debris

5.1.2 Plan for limiting probability, if applicable.

5.2 Assessment of Collisions with Small Debris During Mission Operations

5.2.1 Description of primary mission failure modes from meteoroid or
orbital debris impact

5.2.2 Description of design measures taken to protect against impacts.

6\. DESCRIPTION OF POSTMISSION DISPOSAL PROCEDURES AND SYSTEMS

6.1 Description of Postmission Disposal Option and Disposal System

6.1.1 Statement of disposal option exercised

6.1.2 Disposal plan and description of supporting systems

6.1.3 Source for analysis

6.2 Assessment of Potential Failures that Prevent Successful Postmission
Disposal (N/A)

7\. ASSESSMENT OF SURVIVAL OF DEBRIS FROM THE POSTMISSION DISPOSAL
ATMOSPHERIC REENTRY OPTION

7.1 Verification that Surviving Debris is Within Guidelines

7.2 Source for analysis

**REVISION NOTICE**

Preliminary Release For Comments: May,1997

Added Delta upper stage analysis and integrated figures from NSS 1740.14

Expanded description of observatory and added illustrations.

Transformed to CDR-45 days report format

Initial Issue: July 2 1997

Revision 1: November 1997

Revised Section 2.2.3. Added results of long-term orbit analysis to
Section 6.1.1.

**TABLE of TBRÕs**

Section 7.1: TBR awaiting analysis of Delta 2 second stage reentry
casualty area.

# BRIEF BACKGROUND ON PROGRAM AND PROGRAM MANAGEMENT

## Mission Description

In-situ measurements made for the past 35 years have yielded a wealth of
statistical information about the magnetosphere and its constituent
plasma regimes. They have also provided many examples of dynamic changes
of magnetospheric plasma parameters at specific times and places in
response to changes in the solar-wind input and to internal disturbances
related to substorms. However, statistical global averages and
individual events are not sufficient to understand the dynamics and
interconnections of this highly dynamic system. Fundamental questions
concerning plasma entry into the magnetosphere, global plasma
circulation and energization, and the global response of the
magnetospheric system to internal and external forcing functions remain
unanswered. To further our understanding of the physical processes that
affect the magnetosphere requires the nearly instantaneous measurement
of its topology, that is, magnetospheric imaging.

## Program/Project Objectives

To address the basic questions above, IMAGE will provide imaging of
three general magnetospheric regions: a) the magnetopause, boundary
layer, cusp, and auroral zone; b) the plasmasphere; and c) the inner
plasma sheet, ring current, and trapped radiation. IMAGE will also
provide images of near-Earth interplanetary space. The data acquired in
each of these regions will be used to determine the global structure of
the magnetosphere, characterize the connectivity between magnetospheric
regions, and place results in a global context with previous in-situ
measurements.

The overall objective of IMAGE is to determine the global response of
the magnetosphere to changing conditions in the solar wind.

## Program/Project Schedule

The IMAGE program carried out a successful Mission Requirements Review
in September, 1996. The Confirmation Review allowing the project to
proceed to phase C/D was held February 25-27, 1997. The mission CDR was
held on August 19-20, 1997. Table 1 is a list of scheduled review dates.

Table 1 - Scheduled Review Dates

+------------------------------------+---------------------------------+
| > **[Review]{.underline}**         | > **[Date]{.underline}**        |
+------------------------------------+---------------------------------+
| > MRR                              | > September 10-12, 1996         |
+------------------------------------+---------------------------------+
| > Confirmation Review              | > February 25-27, 1997          |
| >                                  |                                 |
| > (Mission PDR and NAR)            |                                 |
+------------------------------------+---------------------------------+
| > CDR                              | > August 19-20, 1997            |
+------------------------------------+---------------------------------+
| > MOR                              | > June, 1998                    |
+------------------------------------+---------------------------------+
| > PER                              | > August 1999                   |
+------------------------------------+---------------------------------+
| > PSR                              | > October 1999                  |
+------------------------------------+---------------------------------+
| > FRR                              | > December 1999                 |
+------------------------------------+---------------------------------+

Instrument delivery to SwRI for integration into a Payload will be in
January 1999. Integrated and tested Payload delivery to LMMS for
integration into the Spacecraft will be in June 1999. Scheduled launch
date is January 1, 2000. The planned mission duration is two years on
orbit.

## Responsible program or project manager

The Principal Investigator (P. I.), Dr James Burch, is accountable to
NASA for the scientific success of the investigation. In accordance with
NASA's transfer of program management responsibility to its Centers,
Explorer program management responsibility has been assigned to GSFC. In
this role, GSFC is responsible for the exercise of NASA's fiduciary
responsibility to ensure that Explorer missions are achieved in
compliance with committed cost, schedule, performance, reliability, and
safety requirements. The P. I. and his support team at Southwest
Research Institute (SwRI), will be responsible for the overall
management of the investigation. With the assistance from the
Co-Investigator institutions, SwRI will be responsible for the following
elements of the overall program:

a)  Development of science requirements and the science mission plan;

b)  Development of the instrumentation and support systems;

c)  Procurement of the IMAGE spacecraft;

d)  Integration of the spacecraft and science payload;

e)  Support for the integration of the spacecraft to the Delta 7326
    launch vehicle;

f)  Initial on-orbit activation of the spacecraft including instruments;

g)  Training of the GSFC Flight Operations Team;

h)  Development and operation of the IMAGE science operations functions;

i)  Acquisition and validation of the data base and implementation of an
    open data policy for the mission;

j)  Development and execution of a comprehensive education and public
    outreach program featuring magnetospheric science;

k)  Analysis and publication of peer-review science publications;

l)  Presentation of science results to the scientific community and the
    public at large; and

m)  Mission systems engineering and verification.

GSFC will be responsible for the following elements of the IMAGE
program:

a)  Oversight of the MIDEX IMAGE mission on behalf of NASA;

b)  Procurement of the Delta 7326 launch vehicle; and

c)  Development and operation of the Science and Mission Operations
    Center (SMOC).

# DESCRIPTION OF DESIGN AND OPERATIONS FACTORS

## Hardware

### Physical description of main structure

The main body of the IMAGE spacecraft, illustrated in Figure 1 and
Figure 2, consists of a honeycomb aluminum octagonal prism 2.2 meters in
diameter (across the flats, x and y directions) and 1.362 m tall (z
direction). There is an internal octagonal payload deck perpendicular to
the z axis and about midway through the prism. The payload deck supports
the complement of 6 instruments (counting FUV-WIC, -GEO, and -SI as one,
and HENA, MENA, LENA, EUV, and RPI as one each) and the Central
Instrument Data Processor (CIDP). The outer panels have face sheets
approximately 9 mils thick (both sides). There are interior shear panel
structures between the payload deck and the -z panel which support the
spacecraft equipment including the Spacecraft Control Unit (SCU),
battery, Power Distribution Unit, and other support hardware.

There are two axial antenna booms protruding in the +z and -z directions
10 m each. They are composed of fiberglass lattice structures with a
small aluminum end plate, supporting a beryllium copper wire antenna
along their entire length. There are 4 radial antennas, composed of
beryllium copper with beryllium copper end masses, extending 250m each
in the +x, +y, -x, and -y directions.

Cutouts in the side panels permit the imaging instruments to see out in
the plane of rotation of the spacecraft. The six imaging instruments
require apertures in a variety of sizes and shapes, up to approximately
40 x 54 cm. Several of the instruments have aluminum collimators
extending through their cutouts up to a distance of several cm.

The Delta 3rd stage is a STAR 37FM solid rocket motor with nutation
damping and a yo-yo despin system. The Delta 2nd stage is a standard
Delta launch vehicle second stage with a spin table connecting it to the
3rd stage.

The two yo weights released by the Delta third stage each consist of a
flat 978 cm (385) inch by 0.65 cm (0.25 inch) kevlar cable connected to
a 5 x 5 x 7.5 cm (2 x 2 x 3 inch) solid metal weight.

![](media/image2.emf)

Figure 1: IMAGE orientation in Space (Satellite shown over North Pole)

![](media/image3.emf)

Figure 2: Cutaway illustration of IMAGE internal structure

### Description of surfaces/materials exposed to space

The majority of the observatory area exposed to space is composed of
solar cells. The remainder of the structure is aluminum, and there is a
radiator band about the middle of the spacecraft (at the same level as
the payload deck and thermally connected to it) with a Ceramic Optical
Surface Reflector (COSR) surface. Instrument collimators have a variety
of surfaces including optically black paints or coatings to minimize
scattering into their detectors.

The radial stranded Beryllium Copper antennas are not coated and are
exposed to space. The axial fiberglass antenna booms and the Beryllium
Copper wires they support are not coated and are exposed to space.

Exposed surfaces of the Delta third stage include the aluminum payload
support structure, the combustion chamber shell, and the rocket exhaust
nozzle. Exposed surfaces for the Delta second stage include the
second-stage miniskirt and support truss, the spin table, the
pressurized gas spheres (helium and nitrogen), the propellant tanks, and
the rocket exhaust nozzle.

### Description of spacecraft components most sensitive to debris impact

The critical components of the observatory which maintain orientation
include:

a)  A magnetometer and a star tracker;

b)  A magnetic torque bar;

c)  The Spacecraft Control Unit (SCU) for calculation of orientation and
    commanding the torque bar;

d)  The Power Distribution Unit (PDU) to supply power to the above;

e)  A sufficient area of solar panel to activate the above.

There is a battery on the observatory which is used only during eclipse
passage. Because of the very high angular momentum of the observatory,
complete loss of power during these passages would have little or no
adverse effect. The magnetic torquer is not effective except during the
low-altitude part of the orbit (\< 2 Earth radii).

There are several critical components of the Delta second and third
stages which contribute to attitude control and sequencing. These
components are situated as on other Delta flights with similar
configuration. They are only needed during the first hour of the mission
(before spacecraft separation), so their exposure to damage by debris
impact is minimal.

### Description and location of pressurized volumes

There is a sealed 21-Ah Super Ni-Cad battery to provide electrical
power. The FUV GEO sensor contains three 5 ml sealed volumes of Oxygen
at a pressure of 205 torr. The Delta second stage contains pressurized
volumes of nitrogen and helium.

### Description of on-board propellants

There are no propellants on board the observatory. Several of the
instruments are considering the use of "dimple motors" to activate their
cover releases (The covers remain permanently attached, swinging out of
the way on hinges or retracting into cover housings.) Dimple motors are
designed to completely contain the gases produced by their activation.
All dimple motors will be activated during the on-orbit deployment
process.

All of the usable propellant in the Delta third stage is consumed during
its burn. The Delta second stage will have some remaining propellant
after third stage separation. This propellant will be burned off in a
depletion burn after separation from the third stage, which will change
only the inclination of the second stageÕs orbit.

### Description and location of fuel storage and transport systems

There is no fuel storage or transport system on the observatory.

The Delta third stage STAR 37FM is a solid fuel rocket. Its fuel is
stored and used internal to the combustion chamber.

The Delta second stage contains standard facilities for fuel storage and
transport. Its fuel is all expended either during the launch or in the
depletion burn immediately after third stage separation.

### Description of range safety systems

No range safety system is planned for the payload. The Delta third and
second stages may contain range safety systems according to standard
practice.

### Description of systems containing stored kinetic energy

No onboard system contains stored kinetic energy other than the energy
associated with the rotation of the entire spacecraft about the z axis.

## Mission Parameters

### Number of Spacecraft

There is only one spacecraft in the IMAGE mission.

### Launch Date and Time

The launch date is nominally January 1, 2000 and no later than March
2000. The time of day is chosen to satisfy a right ascension of the
ascending node of 0 or 180 degrees from the vernal equinox.

### Mission Orbit

The mission orbit, Figure 1, is an elliptical polar orbit (90.0 degrees
inclination) with a perigee altitude of 1000 km and an apogee altitude
of seven (7) Earth radii. The argument of perigee at launch is 320
degrees, and during the mission it progresses to 220 degrees. During the
mission, apogee is always in the northern hemisphere and is over the
north pole at the mission half way point. The ascending node is chosen
to be either 0 or 180 degrees. Thermal and power considerations
constrain the orbit to be within 15 degrees of one of these target
ascending node locations.

The Delta third stage will occupy the same orbit after separation.

After separation, the Delta second stage will occupy an orbit with a
perigee altitude of 185 km, an apogee altitude of 1073 km, and an
inclination of 89.79 degrees (the inclination may be affected by the
depletion burn).

### Flight Attitude

The spacecraft is spin stabilized with its axis of spin perpendicular to
the orbit plane. The sense of spin is retrograde to the orbit
circulation (i.e. the spin vector is opposite to the orbit angular
momentum vector).

The Delta third stage is not stabilized after de-spin, but its residual
spin and nutation damper will tend to keep its axis aligned in the plane
of the orbit.

The Delta second stage will be stabilized by its on-board control system
until its batteries are discharged, after which it is not stabilized.

![](media/image4.wmf){width="6.083333333333333in" height="5.53125in"}

Figure 3: IMAGE Mission Orbit

#  Assessment of Debris Released During Normal Operations

## Debris Released During Staging, Payload Separation, and Payload Deployment

### Description of debris released, size, mass, area, initial orbit, and lifetime

The Delta second stage is released at separation. The separated second
stage has an overall area of approximately 9.8 m^2^ (105 square feet)
and a Ballistic Coefficient (W/CdA) of 8.4 lb/ft^2^, which gives an area
mass density of approximately 0.024 m^2^/kg. As described above, after
separation, the Delta second stage will occupy an orbit with a perigee
altitude of 185 km, an apogee altitude of 1073 km, and an inclination of
89.79 degrees (the inclination may be affected by the depletion burn).
Orbit lifetime for the 1073x185 km orbit is predicted by the Debris
Assessment Software (Version X.09) to be 0.0638 years.

A system of yo weights with kevlar cables is used to de-spin the Delta
third stage. The yo weights are released to carry away the excess
angular momentum associated with the spin of the stage and observatory.
The yo weights are compact metal masses, with attached flat ribbon
kevlar cables 9.77 meters (385 inches) long. Projected area of the
released yos is approximately 0.0362 m^2^.(56 square inches). Mass of
the yo weight plus the kevlar cable is approximately 0.78 kg (1.72 lbs).
Area-to-mass ratio is approximately 0.046 m^2^/kg. The yo weights and
cables occupy the mission orbit. As described above, the mission orbit
is an elliptical polar orbit (90.0 degrees inclination) with a perigee
altitude of 1000 km and an apogee altitude of seven (7) Earth radii. The
argument of perigee at launch is 320 degrees. Orbit lifetime is not
predicted accurately by either Figure 3-2 of the NASA Safety Standard
Guidelines and Assessment Procedures for Limiting Orbital Debris (NSS
1740.17) or by the Debris Assessment Software (Version X.09).
Extrapolating on Figure 3-2 (reproduced below as Figure 4 for
convenience) indicates that the orbit lifetime may be on the order of
3,000,000 years.

The Delta third stage is the last piece of debris released during launch
operations. The separated third stage has an area of approximately 3.05
m^2^ (32.8 square feet) and a Ballistic Coefficient (W/CdA) of 11.2
lb/ft^2^, which gives an area mass density of approximately 0.02
m^2^/kg. The third stage occupies the same mission orbit as the yo
weights. As with the yo weights, orbit lifetime is not accurately
predicted by either Figure 3-2 of the NASA Safety Standard Guidelines
and Assessment Procedures for Limiting Orbital Debris (NSS 1740.17) or
by the Debris Assessment Software (Version X.09) and may be on the order
of 3,000,000 years.

![](media/image5.wmf){width="6.492361111111111in"
height="6.161111111111111in"}

Figure 4: (NASA NSS 1740.14 Figure 3-2) Orbit dwell times below 2000km
for debris released in high eccentricity orbit. Radiation pressure
effects and lunar and solar gravity perturbations neglected in the
calculation or orbit lifetime.

### Calculated area-time for debris greater than 1 mm (mm^2^-year)

The area of the second stage given above is 9.8 m^2^. The predicted
orbital lifetime given above is 0.0638 years. The area-time product of
the second stage is therefore 0.625 m^2^-years.

The area of the yos is given above as 0.036 m^2^. The orbital lifetime
is taken to be 3,000,000 years in accordance with the fairly uncertain
estimate given above. During each orbit, approximately 2.55% of the
orbit period is spent below 2000 km altitude. Taking this factor times
the estimated orbit lifetime times the area gives an area-mass product
for each yo of 2754 m^2^-years. There are two yos. Using the above
approximations indicates that the total area-time product for the despin
system will be approximately 5508 m^2^-years.

The area of the Delta third stage is given above as 3.05 m^2^. The
orbital lifetime is taken to be 3,000,000 years in accordance with the
estimate above. During each orbit, approximately 2.55% of the orbit
period is spent below 2000 km altitude. Taking this factor times the
estimated orbit lifetime times the area of the third stage gives an
area-mass product of 233,000 m^2^-years. The total object-year product
for the Delta third stage will be approximately 75,000 object-years.

### Calculated object-time for debris greater than 1 mm (mm^2^-year)

The Delta second stage is a single body for which the orbital lifetime
has been calculated to be 0.0638 years. The total object-time product
for the Delta second stage will be approximately 0.0638 object-years.

There are two yo weights in the despin system. The approximations above
indicate that the lifetime of both objects may be approximately
3,000,000 years, of which 2.55% ( 76,500 years) will be spent below an
altitude of 2,000 km. The total object-time product for the despin
system will be approximately 153,000 object-years.

The Delta third stage is a single body. The approximations above
indicate that its lifetime may be approximately 3,000,000 years, of
which 2.55% ( 76,500 years) will be spent below an altitude of 2,000 km.
The total object-time product for the Delta third stage will be
approximately 76,500 object-years.

### Calculated time for removal of debris from GEO altitude to at least 300 km below GEO altitude

The Delta second stage will never attain geosynchronous altitude and
will soon reenter.

The yo weights of the despin system and the third stage will all be in a
highly elliptical orbit whose apogee is above geosynchronous altitude.
Their long-term disposal will be the same as for the Observatory. See
Section 6 for details.

### Source for analysis

All orbital lifetime calculations presented in this section were made by
or interpolated from charts preseted in the NASA Safety Standard
Guidelines and Assessment Procedures for Limiting Orbital Debris (NSS
1740.17) or by the Debris Assessment Software (Version X.09).

## Debris Released During Mission Operations (N/A)

No debris is released during normal or contingency mission operations.

# Assessment of orbital debris generated by explosions and intentional breakups

## Explosions from On-Board Stored Energy

### Description of failure modes leading to explosion

The only credible scenario leading to explosion aboard the Observatory
is mismanagement of battery charging leading to battery explosion. No
other sources of energy capable of causing explosion are available on
the Observatory.

The second stage will contain an undetermined amount of unburned
propellant following its last cutoff during ascent. This propellant
could cause an explosion if allowed to burn in an uncontrolled manner.

### Description of systems involved in explosive failure

The complete battery including pressure vessel and electrolyte has a
mass of 22.5 kg. Its active chemical composition is a Nickel-Cadmium
system. The energy storage capability of the battery is 21 Amp hours at
28 Volts. The overall battery energy density is no more than
21\*3600\*28 Joules/ 22.5 kg = 94,080 J/kg at full charge. Later in the
mission, as the maximum solar array output voltage decreases due to
radiation exposure, the maximum energy density will decrease.

The second stage will contain A-50 propellant and N~2~O~4~ oxidizer. The
effective energy density of the propellant can be estimated from its
specific impulse to be approximately 4.9\*10^6^ Joules/kg. The total
mass of unburned propellant cannot be accurately estimated, but will be
a small fraction of the total available propellant.

### Estimated probability of explosion

Pre-launch testing will mitigate the chance of an error in the design of
the battery or charge control circuitry which could lead to dangerous
conditions. The charge control circuitry is composed of rad-hard
components with sufficient redundancy to render the chance of a latchup
or other radiation-caused event negligible. Neither of these causes is
seen as a credible scenario leading to battery failure. The design of
the spacecraft is such as to protect the battery from debris or meteor
impact by enclosing it within honeycomb of at least 0.020 inches total
thickness. This protection is adequate to make the probability of
impact-caused explosion also negligible. The probability of a battery
explosion is very small and is not quantified in this assessment.

The probability of an explosion in the second stage before its depletion
burn uses up all of its remaining propellant is also very small due to
the short elapsed time before the depletion burn. This probability is
not quantified in this assessment.

### Detailed plan for safing structure after completion of mission

The Delta Second Stage will burn off its remaining propellant in a
depletion burn (directed so as to change only the inclination of its
orbit) after separation from the third stage. The probability of
explosion will drop to zero after completion of this burn.

The battery is charged by solar arrays on the exterior of the
spacecraft. The orbit passes through radiation belts which will be
detrimental to the performance of the solar cells, both in charging
current and in open-circuit voltage produced. Analysis indicates that
within approximately 5 years, the open-circuit voltage produced by the
solar cells will be less than the minimum charging voltage of the
battery. After this happens, the battery will discharge within two
eclipse-season orbits to a zero charge state and will never recover its
charge. The stored charge in the battery is therefore not a credible
source of explosions after the five-year point.

## Intentional Breakups (N/A)

The IMAGE program will not involve any intentional breakup of on-orbit
hardware.

# ASSESSMENT of Debris Generated by On-Orbit Collisions

## Assessment of Collisions with Large Objects During Mission Operations

### Estimated probability of collision with intact space systems or large debris

We compute the estimated probability of the Observatory undergoing
collision in the Table below. The altitude bands are chosen to
accommodate Figure 5-1 of the NASA Safety Standard Guidelines and
Assessment Procedures for Limiting Orbital Debris (NSS 1740.17), which
provides the area flux values. That figure is reproduced below (Figure
5) for convenience.

![](media/image6.wmf){width="6.394444444444445in"
height="5.770833333333333in"}

Figure 5: (NASA NSS 1740.14 Figure 5-1) Cross-sectional area flux of
intact space systems and large orbital debris.

The Total Time in each altitude band is calculated based on the mission
orbit described in paragraph 2.2. Table 2 shows cumulative time spent in
each band during the mission.

Table 2 - Cumulative Time In Each Altitude Band

  ----------------- ----------------- ------------------------- --------------------
  **Altitude Band** **Total Time in   **Area Flux**             **Time-Flux
                    Band**                                      Product**

  **(km)**          **(years)**       **(Impacts/m^2^/year)**   **(Impacts/m^2^)**

  1000-1100         0.007737          2.50E-06                  1.93E-08

  1100-1370         0.007342          1.30E-06                  9.54E-09

  1370-1600         0.004337          1.70E-06                  7.37E-09

  1600-1690         0.001495          8.00E-07                  1.2E-09

  1690-1790         0.001571          4.00E-07                  6.28E-10

  1790-1920         0.00193           2.00E-07                  3.86E-10

  1920-2000         0.001136          1.20E-07                  1.36E-10

                                      **Sum (impacts/m^2^)**    3.9E-8
  ----------------- ----------------- ------------------------- --------------------

Taking the Sum of the Time times Flux products and multiplying by the
projected area of the spacecraft, approximately 6 m^2^, gives the total
probability of collision of **2.3e-7**. This is well below the guideline
of 0.001.

### Plan for limiting probability, if applicable.

Since the probability of collision is lower than the guideline by a
factor of approximately 4000, no action is contemplated to reduce that
probability.

## Assessment of Collisions with Small Debris During Mission Operations

### Description of primary mission failure modes from meteoroid or orbital debris impact

There are several components critical to Observatory operation which are
susceptible to debris impact. They have been divided into two lists in
Table 3. The first list contains components which are exposed to direct
debris impacts. The second list contains components which are protected
inside the aluminum honeycomb body of the Spacecraft. The systems
identified as critical to operation of the science instruments but not
critical to Observatory survival, attitude maintenance, and
communication are marked with an asterisk (\*).

Table 3 - Critical Components Susceptible to Debris Impact

  -----------------------------------------------------------------------
  **Exterior Critical Items**

  Sun Sensor, Star Tracker (One or the other must be operational)

  Solar Panels

  Instrument Collimators/Filters\*

  **Interior Critical Items**

  Power Distribution Unit

  Battery

  Spacecraft Control Unit

  Spacecraft Cable Harnesses

  Solid State Memory\*

  Central Instrument Data Processor\*

  Payload and S/C to P/L Cable Harnesses\*
  -----------------------------------------------------------------------

Destruction of or damage to any of the instrument exterior components
would adversely affect operation of that instrument without affecting
the remainder of the payload. Damage to the solar panels may adversely
affect the power cycle of the observatory during eclipse season or, if
severe enough, during all operations. Damage to both the sun sensor and
the star tracker would disable all baseline methods of attitude
determination. It is possible that one or more instruments could be used
as an attitude reference, but this possibility has not been studied in
depth.

Damage to any of the components on the interior critical items list
would effectively terminate science operations.

Postmission disposal plans would not be affected by damage to any
component of the observatory.

### Description of design measures taken to protect against impacts.

The design of the Observatory, with exterior structural panels
protecting most of the components, minimizes the probability of
destructive collision with debris. All sensitive exterior components are
required to be exterior components to fulfill their design function;
shielding of these components is not practical. All interior components
are protected from impacts with objects up to at least 0.020 inches in
diameter by the exterior panels. The costs of adding additional
shielding (in terms of mass and budget) are evaluated to be higher than
the benefits associated with the reduced probability of damage.

The post-mission disposal plans do not require functionality of any
components of the IMAGE Observatory. Therefore, the probability of
collision with small debris preventing postmission disposal according to
plan is zero.

# Description of Postmission Disposal Procedures and Systems

## Description of Postmission Disposal Option and Disposal System

### Statement of disposal option exercised

The planned disposal procedure for the IMAGE observatory is deactivation
in mission orbit. The yo weights and Delta third stage will also remain
in this orbit. No engines or other devices are used to alter the orbit.
The science orbit for this mission is used very little by other
spacecraft. Table 4 shows the time fraction of an orbit which is spent
in any of the sensitive regions as defined in Figure 6-1 of the NASA
Safety Standard Guidelines and Assessment Procedures for Limiting
Orbital Debris (NSS 1740.17).

Table 4- Fraction of Time Spent in Sensitive Regions

  ---------------------------------------------------- ------------------
  **Sensitive Orbit Regions**                          **Fraction of
                                                       time**

  Low Earth Orbit (\<2,500 km altitude)                3.2%

  Semi-Synchronous orbit (19,900 km \< altitude \<     0.8%
  20,500 km)                                           

  Geo-Synchronous orbit (35,288 km \< altitude \<      2.1% \*\*
  36,088 km)                                           
  ---------------------------------------------------- ------------------

\*\*The polar inclination of the orbit combined with the precession of
perigee at a rate of approximately 50 degrees per year means that the
orbit will pass through the sensitive equatorial geo-synchronous region
only 4 times every 7 years. The 2.1% figure is thus an extremely
pessimistic upper limit on the time spent in the geosynchronous
equatorial region.

The long-term effect of luni-solar perturbations to the orbit has been
analyzed by Nicholas L. Johnson of Johnson Space Flight Center, SN3. The
primary effect predicted is an oscillation in the perigee altitude of
objects in the mission orbit. The perigee will oscillate between values
of approximately 500 km and approximately 1600 km, with a period of
oscillation of approximately 1400 days. (The initial direction of change
will be toward a high-altitude perigee, so there are no adverse effects
on the science mission.) No other significant effects on the orbit were
reported. Because of their cyclical nature and small amplitude, these
perturbations do not substantially affect any of the calculations
presented in this document.

### Disposal plan and description of supporting systems

After separation, the inert (spent) Delta third stage and yo masses will
serve no further mission functions. No further operations involving them
are planned.

At termination of operations for the Observatory, the instruments will
be commanded to a safe condition to prevent sun damage to the sensors.
The attitude control system will be commanded to maintain the
Observatory in the science mission attitude. The system will be placed
in a state where removal and re-application of battery power will cause
it to come up in a stable condition, if that is not the normal flight
mode. The solid state recorder will continue to operate but will
over-write data older than approximately 21 hours (and no science data
will be available except from instruments which are always sun-safe,
such as RPI). Broadcast transmission of real-time data for NOAA or other
users will continue unless commanded off to clear radio bandwidth. The
Observatory can be reactivated from this state by uploading appropriate
ephemeris data, checking operation of the attitude control system, and
reactivating the instrument. Operations aboard the Observatory will
continue, cycling the battery during eclipse season.

As solar panel output deteriorates, the Spacecraft computer will
gracefully load-shed by powering off instruments and nonessential
systems. Within about five years, radiation damage to the solar panels
will cause their open-circuit voltage output to drop below the charging
voltage for the battery. The next one or two eclipses will discharge the
battery to its full-discharge state and it will not recover. The
Observatory will power down during and recover after each succeeding
eclipse. At some later time the panel output will drop to a voltage too
low to power up the Spacecraft Control Unit. At this point the
Observatory will no longer be operable.

### Source for analysis

All analysis used in this section was taken from the NASA Safety
Standard Guidelines and Assessment Procedures for Limiting Orbital
Debris (NSS 1740.17) or done by the Debris Assessment Software (Version
X.09).

## Assessment of Potential Failures that Prevent Successful Postmission Disposal (N/A)

No obstacles to the postmission disposal plans described above have been
identified.

# Assessment of Survival of Debris from the Postmission Disposal Atmospheric Reentry Option

## Verification that Surviving Debris is Within Guidelines

Atmospheric re-entry is not contemplated as a disposal option for the
IMAGE observatory, Delta third stage, or de-spin yo masses.

The Delta second stage will reenter less than one month after
separation. Analysis of debris survival of reentry is not complete at
this point, but should not differ substantially from normal Delta second
stage reentry (**TBR**). A complete analysis or reference to a
previously completed analysis will be appended to this Debris Assessment
Report as soon as possible.

## Source for analysis

All analysis used in this section will be taken from the NASA Safety
Standard Guidelines and Assessment Procedures for Limiting Orbital
Debris (NSS 1740.17) or done by the Debris Assessment Software (Version
X.09).