Collision with orbital debris is a hazard of growing concern as historically accepted practices and procedures have allowed man-made objects to accumulate in orbit. To limit future debris generation, NASA Management Instruction (NMI) 1700.8, "Policy to Limit Orbital Debris Generation," was issued in April of 1993. The NMI requires each program to conduct a formal assessment of the potential to generate orbital debris. This standard serves as a companion to NMI 1700.8 and provides each NASA program with specific guidelines and assessment methods to assure compliance with the NMI.

Each main debris assessment issue (e.g., Post Mission Disposal) is developed in a separate chapter. For the reader who needs just an overview of the debris issues, consult the guidelines descriptions in chapter 2.

The standard was developed jointly by the Office of Safety and Mission Assurance (Code Q) and the Johnson Space Center Space Physics Branch. Comments, questions, or suggestions concerning this document should be directed to Code QS.

Frederick D. Gregory
Associate Administrator for
Safety and Mission Assurance

TABLE OF CONTENTS

CHAPTER PAGE

PREFACE i

1. INTRODUCTION 1-1

1.1 Scope and Purpose 1-1

1.2 Overview of NASA Management Instruction 1700.8 1-1

1.3 Objectives of Debris Assessments 1-1

2. CONDUCTING THE DEBRIS ASSESSMENT: AN OVERVIEW 2-1

2.1 Structure of the Guidelines Document 2-1

2.2 Performing Debris Assessments 2-2

3. ASSESSMENT OF DEBRIS RELEASED DURING NORMAL OPERATIONS 3-1

4. ASSESSMENT OF DEBRIS GENERATED BY EXPLOSIONS AND INTENTIONAL BREAKUPS 4-1

4.1 Accidental Explosions 4-1

4.2 Intentional Breakups 4-2

5. ASSESSMENT OF DEBRIS GENERATED BY ON-ORBIT COLLISIONS 5-1

6. POSTMISSION DISPOSAL OF SPACE STRUCTURES 6-1

7. SURVIVAL OF REENTERING SPACE SYSTEM COMPONENTS AFTER POSTMISSION DISPOSAL 7-1

8. FORMAT FOR ASSESSMENT REPORTS 8-1

8.1 Format for Report Issued at PDR 8-1

8.2 Format for Report Issued Prior to CDR 8-3

APPENDICES PAGE

APPENDIX A: DEFINITION OF TERMS A-1

APPENDIX B: BACKGROUND ON ORBITAL DEBRIS B-1

TABLES

TABLE PAGE

2-1 Debris Assessment Issues and Corresponding Guideline Descriptions 2-2

5-1 L Factor Values for Critical Surfaces on Vehicles Stabilized

Relative to the Velocity Vector 5-6

7-1 Selected Properties for Materials Commonly Used in Spacecraft Fabrication 7-5

FIGURES

FIGURE PAGE

3-1 Orbit lifetime for debris released in low altitude, low eccentricity orbits 3-7

3-2 Orbit dwell time below 2000 km for debris released in high eccentricity orbit 3-8

3-3 Initial orbit apogee and perigee altitude required to lower debris orbit to

300 km below GEO altitude in 25 years 3-9

5-1 Cross-sectional area flux of intact space systems and large orbital debris 5-7

5-2 Cross-sectional area flux for orbital debris as a function of debris diameter

for spacecraft in low Earth orbit 5-8

5-3 Cross-sectional area flux for meteoroids as a function of meteoroid diameter 5-9

6-1 Disposal regions and storage orbit options for postmission disposal 6-2

6-2 Area-to-mass contours for 25-year orbit lifetime for space systems left in

low altitude, low eccentricity orbit 6-8

6-3 Area-to-mass contours for 25-year orbit lifetime for space systems left in

highly eccentric orbit 6-9

7-1 Flow chart for calculating total debris casualty area for uncontrolled reentry 7-3

1.1 SCOPE AND PURPOSE

This document serves as a companion to NASA Management Instruction (NMI) 1700.8, and provides specific guidelines and methods to comply with the NASA policy to limit orbital debris generation. The guidelines serve to help ensure that launch vehicles, upper stages, and payloads meet acceptable standards for limiting orbital debris generation. This document should be used by the program manager or project manager as the primary reference in conducting debris assessments. The standard establishes guidelines and provides supporting analysis tools for: (1) limiting the generation of orbital debris, (2) assessing the risk of collision with existing space debris, and (3) assessing the potential of spacecraft-generated debris fragments to impact the Earth's surface. In addition to guidelines and methods for assessment, this volume provides formats for the debris assessment reports. Two appendices are used to define frequently used terms and to provide summary background information.

Another document, entitled "Reference Manual for Orbital Debris Assessments," provides more in-depth background and technical information. In addition, debris assessment software is available to support the assessment of particular guidelines and to evaluate mitigation measures.

1.2 OVERVIEW OF NASA MANAGEMENT INSTRUCTION 1700.8

NASA Management Instruction (NMI) 1700.8 states "NASA's policy is to employ design and operations practices that limit the generation of orbital debris, consistent with mission requirements and cost-effectiveness." The NMI requires that each program or project conduct a formal assessment for the potential to generate orbital debris.

The debris assessment must address the potential for orbital debris generation that results from normal operations and malfunction conditions, and on-orbit collisions. The assessment must also address provisions for postmission disposal. Malfunction conditions refer to those credible failure scenarios or conditions that can result in the direct generation of orbital debris or that can disable the spacecraft to preclude postmission disposal. Examples of orbital debris generated during normal operations include items such as lens covers, shrouds, and staging components that are released into the environment. An on-orbit explosion is an example of debris generation by malfunction. Examples of debris generation by collisions include immediate debris generation by collisions with large objects and by loss of control of a spacecraft or payload as a result of impact with small debris during mission operations.

To satisfy the NMI, the program or project manager may need to plan for such things as:

1.3 OBJECTIVES OF DEBRIS ASSESSMENTS

Each program or project should attempt to meet all pertinent guidelines. It is understood, however, that satisfying these guidelines must be balanced with the necessity to meet mission requirements and to control costs. If a guideline cannot be met because of over-riding conflict with mission requirements or prohibitive cost impact, this should be specifically noted in the assessment with rationale and justification provided.

As a matter of practice, it is desirable for the program or project to work with the Office of Safety and Mission Assurance during the assessment process. Ideally, the program or project should also use the expertise at NASA centers. The Office of Safety and Mission Assurance at each center can direct programs/projects to groups that can provide assistance with debris assessments. These groups have resources for analyzing complex debris problems which may not be covered in the detail necessary in this standard or in the debris assessment software.

This section provides an overview of what should be covered by a debris assessment. The detailed guidelines and evaluation methods for specific assessment issues are presented in chapters 3 through 7.

As required by NMI 1700.8, "Policy to Limit Orbital Debris Generation," the debris assessment covers two broad areas: the potential for generating debris during normal operations or malfunction conditions, and the potential for generating debris by collision with space debris (natural or human-generated) or orbiting space systems. These two broad areas are broken down into five issues to be addressed in the assessment:

• Debris released during normal operations

• Debris generated by explosions and intentional breakups

• Debris generated by on-orbit collisions during mission operations

• Safe disposal of space systems after mission completion

• Structural components impacting the Earth following postmission disposal by atmospheric reentry

The assessment will be organized around these issues with specific guidelines associated with each. The objective for each program or project is to assess whether all applicable guidelines have been met.

2.1 STRUCTURE OF THE GUIDELINES DOCUMENT

Each of the five assessment issues is covered by a chapter in this guidelines document. Each issue is addressed by specific "guideline" statements. Each guideline statement appears in a bold box at the beginning of a chapter or major section of a chapter.

After the box containing the guidelines, the following sections are presented:

These sections provide information on the reasoning behind the guidelines, a recommended approach for assessing whether a program has met a given guideline, and what kinds of modifications in design or procedures might be considered to bring a program within guidelines. An overview of assessment issues and associated guidelines is presented in table 2-1.

Mitigation measures, analysis support procedures, and technical background are presented in more detail in a companion document, "Reference Manual for Orbital Debris Assessments." Volume I of this document is "Assessment of Debris Mitigation Procedures." Additional technical information is presented in Volume II, "Technical Background for Assessing Orbital Debris Risk." Each volume is organized around the guideline areas used in this standard.

2.2 PERFORMING DEBRIS ASSESSMENTS

Each program should address the applicable guidelines in each of the four areas of normal operations, accidental explosions or intentional breakups, debris collision, and postmission disposal. If atmospheric reentry is used as the postmission disposal option (Section 6), the program should address the guideline on reentry risk contained in Section 7. The guidance provided in the sections titled "Method to Assess

____________________________________________________________________________________

____________________________________________________________________________________

Compliance with the Guidelines" should be sufficient to carry out the assessment. The models in the debris assessment software support the approach and techniques described in this section. Each program is free to use alternative methods or models that they feel are more suitable for their particular program. If alternative methods or models are used, the program should document such methods or models in the debris assessment report.

Two assessment reports should be completed. The first is prepared at PDR and the second 45 days prior to CDR. The PDR debris assessment should identify debris generation issues and , where possible, assess those issues. The CDR assessment should identify, assess, and resolve all debris issues in detail. Chapter 8 of this document provides the specific information that should be included in a debris assessment report.

Although assessments are prepared only at PDR and prior to CDR, it is advisable for each program or project to consider potential debris issues during concept development (Phase A) and development of preliminary requirements, specifications, and designs (Phase B) to estimate and minimize potential cost impacts.

Debris is often released as an incidental part of normal space operations; this type of debris is referred to as operational debris. Since the release can be planned, it can be done in a manner that does not pose a significant risk to other users of space. Small debris—1 mm in diameter (about 1 mg) and larger for LEO, or 5 cm (about 100 gm) and larger for GEO—is a source of concern because it has enough energy to critically damage an operating spacecraft. Larger debris might collide with other large objects in the environment and create clouds of secondary debris fragments.

The probability of a collision occurring with debris released during normal operations depends on the number and size of the debris and on length of time it remains in orbit. The guidelines, therefore, limit the total number of such debris and their orbit lifetimes. Debris released during normal operations includes debris released during staging and payload separation, deployment, and mission operations; tethers and tether fragments left in orbit at the end of mission are also considered operational debris. Upper stages, payloads, and solid rocket motor debris are not covered by these guidelines.

For guideline 3-1a, the value of 0.1 m²-yr is based on the principle that large debris released during normal operations should represent a much smaller risk of collision with other large objects in orbit than do operating spacecraft. Historically, spacecraft have had an average cross-sectional area of about 10 m² and an operational lifetime of three years, giving the average operating spacecraft a 30 m²-yr area-time product. Adopting a guideline value between 0.1 and 1.0 m²-yr would therefore yield a probability of collision that is ~1/30 to 1/300 the collision probability represented by an average operating spacecraft.. A value of 0.1 is used based on mass considerations. Typical operational debris of mass 1 kg and larger, which is large enough to cause complete collisional fragmentation of intact payloads or upper stages, will have an orbit lifetime ~10 years under the 0.1 m²-yr guideline value. If a 1.0 m²-yr value was chosen, such operational debris would have a lifetime of ~100 years, and therefore accumulate in orbit.

Guideline 3-1b limits the total number of debris objects released based on their orbit lifetimes. Based on historical precedent and practice, an acceptable level of risk for released debris damaging another operational spacecraft is <10-6. The guideline value of 100 object-yr was chosen because debris released during normal operations following this guideline will have probability on the order of 10-6 of hitting and potentially damaging an average operating spacecraft.

Tethers present a much greater risk to operating spacecraft than would be expected from their mass and cross-sectional area, and may have a high probability of being severed and thus left in the environment. Consequently, tethers or tether fragments left in orbit after completion of mission require special consideration for the risk they pose to operating spacecraft.

Debris that is not removed from GEO altitude by energy losses resulting from atmospheric drag, a process requiring a low perigee altitude, will remain in the GEO environment for many thousands of years. Therefore, guideline 3-2 limits the accumulation of debris at GEO altitudes and will prevent the development of a significant debris environment, as currently exists in LEO.

For each of the debris objects released during normal operations, calculate the cross-sectional area and orbit lifetime. The total area-time product (guideline 3-1a) will be

å (cross-sectional area ΄ orbit dwell time below 2000 km)

debris larger

than 1 mm

"Orbit dwell time below 2000 km" is defined as the total time spent by an orbiting object below an altitude of 2000 km during its orbit lifetime. If the debris is in an orbit with apogee altitude below 2000 km, the orbit dwell time equals the orbit lifetime.

The total object-time product (guideline 3-1b) will be

å (orbit dwell time below 2000 km)

debris larger

than 1 mm

The following procedure is used to determine orbit dwell time:

100 ΄ (the area-to-mass ratio for the debris object)

The debris assessment software can also be used to perform steps 1-3.

For tethers, which are smaller in two dimensions but orders of magnitude larger in the third dimension than normal space structures, the potential to damage operating spacecraft is much larger than would be expected from the tether mass and cross-sectional area. Consequently, programs using tethers must take extra measures to control the potential for damaging other systems. To limit this type of risk to other users of space, tethers left in orbit after completion of mission or tether fragments created when meteoroids or orbital debris severe the tether are considered operational debris.

To limit the risk presented by the tether debris to operating spacecraft, the maximum tether debris length will be limited by its orbit lifetime. The lifetime must take into account any associated end-point vehicles attached to the tether. The relationship between the maximum allowed tether debris length, , measured in kilometers, and orbit lifetime, T, measured in years, is

[km] = 1 / T [yr]

The orbit lifetime of the tether system will be determined following the same procedure as for evaluation of operational debris for guideline 3-1. In calculating the area-to-mass ratio for orbit lifetime calculations, the average cross-sectional area of the tether system will be the cross-sectional area of the tether measured along its length, (calculated in step 1 below), plus the cross-sectional area of any attached end-point vehicle(s). The mass of the tether system will either be the tether mass, if the tether is detached from the end-point vehicle(s), or the tether mass plus any attached end-point vehicle(s) mass. End-point vehicle mass and cross-sectional area are considered only if the mission plan calls for the tether to remain attached to the end-point vehicle(s). The effective altitude for the system for orbit lifetime calculations will be the altitude of the midpoint of the tether if it is not attached to an end-point vehicle or the altitude of the center of mass for the system if the tether remains attached to an end-point vehicle or vehicles.

The length of tether remaining in the environment after the end of mission, which will be the length of the tether if the mission plan is to leave the tether in the environment at the end of mission, or the length of tether that is expected to be cut off by meteoroid or orbital debris impact during the mission if the mission plan is to retract the tether at the end of mission, must be no larger than .

The length of tether cut during the mission is related to the probability of the tether being cut. To calculate the probability of the tether being cut:

[m²]

[m]

To satisfy guideline 3-1, L may not exceed , if the mission plan is to abandon the tether in orbit; may not exceed if the mission plan is to retract the tether at the end of mission.

This assessment determines the maximum perigee altitude for a debris object if it is to have an apogee altitude at least 300 km below GEO altitude after 25 years as stated in guideline 3- For this analysis use the following steps:

If a program or project does not fall within guidelines, there are a number of mitigation measures that may be taken. These include:

Measures to reduce orbit lifetimes are discussed in "Assessment of Debris Mitigation Measures,"

Volume I of the Reference Manual for Orbital Debris Assessments.

Explosions have been the primary contributor to the orbital debris environment. Some explosions have been accidental with on-board energy sources providing the energy. However, some intentional breakups have occurred as tests or as a means of disposing of spacecraft.

4.1 ACCIDENTAL EXPLOSIONS

Accidental explosions of spent upper stages have been the primary source of debris in LEO. The source of energy for these events has been the structural failure of pressurized volumes or a failure allowing residual hypergolic bipropellant fuels to mix and ignite. The Delta second stage was a source of such debris. Investigations determined that the cause of the explosion was failure of the common bulkhead separating the two fuel components. This failure allowed residual bipropellants to combine, producing an explosion, the most recent of which occurred after 16 years in orbit. The Delta Project Office changed operating procedures to vent the fuels after completion of missions and no stage which has been vented has exploded. In 1978, a Soviet EKRAN satellite in GEO experienced an explosion, with an over-pressurized battery as the suspected cause. After breakup of the Ariane Spot -1 third stage in Sun-synchronous orbit, the Ariane introduced design modifications to prevent future explosions. In cases where design and operations modifications have been made to remove stored energy sources, accidental explosions have been prevented.

On-board energy sources include chemical energy in the form of fuels and explosives associated with range safety systems, energy in the form of pressurized volumes (as in sealed batteries and thermal control, attitude control, or propulsion systems) and kinetic energy (as with control moment gyroscopes).

Concerning the note to guideline 4-1: by keeping the probability of accidental explosion less than 0.0001, the average probability of an operating spacecraft colliding with an explosion fragment larger than 1 mm from that space system will be less than 10-6 per "average spacecraft". An average spacecraft is a spacecraft of average size with average mission lifetime in circular orbit at an altitude through which explosion debris fragments of size 1mm or larger would pass if an explosion occurred. The average probability of collision is the probability of collision averaged over the altitude that would be covered by the breakup cloud.

Under guideline 4-2, explosions caused by stored energy in the form of fluids such as volatile liquids, bipropellant fuels, or electrolytes in batteries will be prevented by venting or using other depletion procedures. In the past, failure to remove energy sources has resulted in explosions occurring anywhere from hours to years after completion of the mission.

When the space system design is being evaluated by a failure mode and effects analysis or some equivalent analysis, the program or project will identify credible failure modes that lead to accidental explosion and estimate the probability of those failure modes occurring. The risk to operating spacecraft caused by an accidental explosion is discussed in "Technical Background for Assessing Orbital Debris Risk," Volume II of the Reference Manual for Orbital Debris Assessments.

During space system design and development, the program will identify sources or potential sources of stored energy and develop and implement a plan for eliminating these sources at the end of mission operations. Safing procedures to be considered might include

4.2 INTENTIONAL BREAKUPS

Intentional breakups have been used to provide safe reentry of space structures and to conduct on-orbit tests. An understanding of the approach taken in the evaluation for intentional breakups requires an understanding of the development of a debris cloud after breakup.

Immediately after breakup, the debris cloud exhibits large spatial and temporal changes in the concentration of the debris. For example, at the point where the breakup occurred there will be no debris at times, while at other times the debris cloud densities will be orders of magnitude above the background. An operating spacecraft may have a small probability of colliding with the debris if the interaction were to occur randomly, but a high probability of collision if it passes through a region of high density concentration. The test program can avoid having such high risk interactions by controlling the time of the test. However, because of the many perturbations that occur to objects in orbit and the sensitivity to the exact time and location of the breakup event, whether operating spacecraft will pass through regions of high debris density concentration can be determined accurately only a few days before the test. Consequently, the assessment and control of this risk must be performed immediately before the test.

Within a few days after breakup, the debris becomes more uniformly distributed within the cloud and the cloud reaches a state called the pseudo-torus. Within a few weeks to a few months after the test, the debris cloud evolves to a shell configuration. By the time the debris cloud reaches the pseudo-torus state, the probability of collision between the debris cloud and other objects in space can be calculated assuming random encounters. This means that the risk to other users of space can be characterized by quantities such as the area-time and object-time product for the debris cloud. These quantities, which depend on the breakup altitude and the general characteristics of the breakup process, can be calculated early in the development process for the testing program.

These guidelines reflect the approach taken within the U.S. space program to limit the debris contribution from on-orbit tests. After the P-78 (SOLWIND) ASAT test, subsequent tests, such as the Delta-180, were reviewed by a safety panel for their near-term threat to operating spacecraft (guideline 4-4) and for their long-term contribution to the orbital debris environment (guideline 4-3).

Debris from intentional breakup released under guideline 4-3 of this section would be no more of a contributor to the long-term growth of the orbital debris environment than debris released under guidelines for normal operations (guideline 3-1). The limit of 1 year for orbit lifetimes for debris larger than 1 mm prevents the accumulation of debris from intentional breakups.

The risk to other users from concentrations within the debris cloud which occur immediately after breakup is limited by guideline 4-4 to no more than the risk represented by other debris deposition events such as release of operational debris.

Guideline 4-5 ensures that all debris from planned satellite destruction as a part of the reentry disposal procedure will deorbit within a few hours.

The evaluation procedure for planned test breakups uses guideline 4-3 for long-term planning conducted during program development, and uses guideline 4-4 for near-term planning conducted immediately (a few days) before the test. The objective of the long-term plan is to understand and control the impact of the test on the space environment in general; that of the near-term plan is to control the risk of damage to operating spacecraft.

The steps for performing the evaluation are

Planned Destruction as a Reentry Procedure (Guideline 4-5)

If the guideline for the breakup altitude of 90 km or below is followed, no additional assessment is required. If the breakup altitude is above 90 km, the disposal procedure is treated as if it were a test breakup and the assessment procedure for on-orbit tests is followed.

This section covers debris generation by random on-orbit collision during mission operations. This includes both the direct generation of debris by collision between the space vehicle and another large object in orbit and the indirect or potential generation of debris when collision with small debris damages the vehicle to prevent its disposal at the end of mission operations, and making it more likely that the vehicle will be fragmented in a subsequent collision with another large object in orbit.

While it remains intact, a spacecraft or upper stage represents a small collision risk to other users of space; however, once it is fragmented by collision, the collision fragments present a risk to other users that is orders of magnitude larger. Because of the large collision velocities, a debris object much smaller than the spacecraft will cause fragmentation (referred to as catastrophic collision). For purposes of evaluation, debris of diameter 10 cm and larger will be assumed to cause such catastrophic collision.

Catastrophic collision during mission operations represents a direct source of debris, and the probability of this occurring is addressed by guideline 5-1. However, if a spacecraft or upper stage fails to perform postmission disposal it becomes a potential source of debris because a structure that is abandoned in orbit may subsequently experience catastrophic breakup. The probability of such an event occurring as a result of damaging impact with small debris is addressed by guideline 5-2.

Guideline 5-1 limits the amount of debris that will be created by collisions between spacecraft or upper stages in LEO or GTO (geosynchronous transfer orbit) and other large objects in orbit. By keeping the probability of collision between a spacecraft or upper stage and other large objects to less than 0.001, the average probability of an operating spacecraft colliding with collision fragments larger than 1 mm from that spacecraft or upper stage will be less than 10-6 per "average spacecraft". An average spacecraft is a spacecraft of average size with average mission lifetime in circular orbit at an altitude through which the fragments from such a collision would pass if the collision occurred. The average collision probability is the probability of collision averaged over the altitude that would be covered by the breakup cloud.

Guideline 5-2 limits the probability of spacecraft and upper stages being disabled and left in orbit at the end of mission, which would contribute to the long-term growth of the orbital debris environment by subsequent collisional fragmentation.

For missions in or passing through low Earth orbit (LEO), the probability of a space system being hit by an intact structure or large debris object during its mission life, P, can be approximated by^*

P = F ΄ A ΄ T (5-1)

where

The orbital debris flux is taken to be 0 for altitudes above 2000 km. The flux for meteoroids 10 cm in diameter or larger is negligible and can be ignored.

For a circular mission orbit the debris flux is taken from figure 5-1 at the mission orbit altitude.

For an eccentric mission orbit, break the orbit into altitude intervals such that the debris flux does not vary by more than a factor of 2 over any interval. Weight the flux in each altitude interval, taken from figure 5-1, by the fraction of time spent by the vehicle in that altitude interval to yield a time-weighted debris flux. The fraction of time spent in an altitude interval is approximated by the ratio of the size of the altitude interval to the difference between the apogee and perigee altitude of the orbit. For eccentric orbits, F is the sum of the time-weighted debris fluxes. The debris assessment software may be used for a more exact calculation of P or for an exact calculation of fraction of time spent in an altitude interval.

The average cross-sectional area is the cross-sectional area averaged over aspect. For a simple convex space system, it is 1/4 the surface area. A simple convex spacecraft body with solar panel wings may be given an average cross-sectional area that is 1/4 the surface area of the spacecraft body plus solar panels.

_______________

* The exact expression for this probability is , which is approximated by Eqn.

5-1 when the product F´ A´ T is less than 0.1

For highly irregular spacecraft shapes, an estimate of the average cross-sectional area may be obtained as follows: determine the view, V, that yields the maximum cross-sectional area and denote the cross-sectional area as Amax . Let A1 and A2 be the cross-sectional areas for the two viewing directions orthogonal to V. Then define the average cross-sectional area as ( Amax + A1 + A2 ) / 2.

For systems in or passing through geosynchronous orbit (GEO), no assessment is required for collision with intact objects or large debris. Collisions between operating spacecraft will be avoided by preventing radio frequency interference between satellites, which follows the historical practice for controlling operating spacecraft in GEO. The population of large debris objects in GEO is thought to present negligible risk to operating spacecraft and can be ignored.

Impact with small (millimeter to centimeter or milligram to gram) meteoroids or debris can cause considerable damage because the impacts usually occur at high velocity (~10 km/sec for debris, ~17 km/sec for meteoroids). An obvious failure mode caused by debris or meteoroid impact is for the impact to puncture a hole in some fluid container, causing leakage. However, many failure modes are not so obvious. For example, the shock pressures produced by an impact on the wall of a pressurized tank can damage the tank wall and cause the tank to rupture even though the impactor may not be large enough to puncture the wall. Some electrical component failures have been suspected to have been caused by debris impacts, either directly by severing wires or indirectly by causing electrical components to short circuit when exposed to melted aluminum ejecta or the plasma cloud generated by a nearby hypervelocity impact.

The following evaluation process is used to determine whether damaging impacts with small debris could reasonably prevent successful postmission disposal. The procedure estimates the probability that meteoroid or orbital debris impacts will cause components critical to postmission disposal to fail. If this estimate shows that there is a significant probability of failure, a full penetration analysis should be conducted to guide any redesign and to validate any shielding design. The procedure outlined below should not be used to design shielding.

To estimate the probability that impacts with small meteoroids or orbital debris will prevent postmission disposal:

s [gm/cm²] = r [gm/cm³] ´ d [cm]

[cm] = K ´ [gm/cm²] , where K=0.07 (5-2)

(5-3)

Surface^$ Front Side Top Bottom Rear

Debris (L_MAN) 3 3 0.01 0.01 0.02

Meteoroids (L_MET) 2 1 2 1 0.2

(5-4)

(5-5)

1. If a LEO program or project has a high probability of colliding with large objects during its mission

life, there are several mitigation measures that may be taken. These include

 

Figure 5-3. Cross-sectional area flux for meteoroids as a function of meteoroid diameter.

6. POSTMISSION DISPOSAL OF SPACE STRUCTURES

The historical practice of abandoning spacecraft and upper stages at the end of mission life has allowed roughly 2 million kg of debris to accumulate in orbit. If this practice continues, collisions between these objects will, within the next 50 years, become a major source of small debris, posing a threat to space operations that is virtually impossible to control. The most effective means for preventing future collisions is to require that all spacecraft and upper stages be removed from the environment in a timely manner. Such a requirement, however, would entail great cost in many cases, and there are regions of space where, for the immediate future, disposal of these systems could be made without creating a significant risk to future users. As a result, a variety of disposal options are presented in the guidelines. These guidelines represent an effective method for controlling growth of the environment while limiting the cost impact on future programs.

To provide a context for the guidelines, three high-value regions of space can be identified. These are

In general, the postmission disposal options are (1) direct retrieval and deorbit, (2) maneuver to an orbit for which atmospheric drag will remove the structure within 25 years, and (3) maneuver to one of a set of disposal regions in which the structures will not interfere with future space operations. Storage orbits in these disposal regions may be used to dispose of space systems at end of mission. These options are summarized in figure 6-1. Most GEO programs will transfer to the super-GEO storage orbit; highly eccentric, high perigee altitude programs might transfer to a sub-GEO storage orbit rather than lowering perigee to reenter. Programs using semisynchronous orbits might use either the low- or high-altitude storage orbit; LEO programs with mission orbit altitudes above 1500 km might choose to transfer to the low-altitude storage orbit rather than transfer to an orbit with a 25-year lifetime. For a program which places a structure in an orbit for eventual atmospheric reentry, there may be restrictions on the disposal maneuver if a significant amount of structure might survive uncontrolled reentry.

The intent of guideline 6-1a is to remove spacecraft and upper stages in LEO from the environment in a reasonable period of time. The 25-year removal time from LEO prevents the debris environment from growing over the next 100 years while limiting the cost burden to LEO programs. Spacecraft and upper stages in mission orbits with perigee altitudes below 600 km will usually have orbit lifetimes less than 25 years and will, therefore, automatically satisfy this guideline. This guideline will have the greatest impact on programs with mission orbit perigee altitudes above 700 km, where objects may remain in orbit hundreds of years if abandoned at the end of mission life.

Guideline 6-1a emphasizes the limitations of using drag enhancement devices to reduce orbit lifetime. Drag enhancement will increase the total area of the spacecraft or upper stage and may do little to reduce the probability of hitting large objects in the environment even though the orbit lifetime is reduced. It is, therefore, essential to demonstrate that drag enhancement does not in fact represent an increased risk to other users of space.

As stated in guidelines 6-1b and 6-2b, disposal orbits between LEO and GEO must have perigee altitudes above 2500 km. Objects in these orbits will have a low probability of collision (the current rate is less than 1 per 1000 years). If a collision does occur, very little debris from that collision will come low enough to place spacecraft in LEO at risk. Depending on the number, size, and orbit characteristics of objects using this disposal orbit option, the separation from LEO may need to be increased in the future.

If spacecraft and upper stages are placed in disposal orbits between LEO and GEO, the constraint on apogee altitude in guidelines 6-1b and 6-2b to be no higher than 500 km below GEO altitude prevents lunar and solar perturbations from causing these structures to interfere with GEO satellite operations. If a collision does occur in these disposal orbits, very little debris from that collision will come high enough to place GEO spacecraft at risk.

In guideline 6-1c only 10 years is allowed for planned retrieval after completion of the mission. This time is shorter than the 25 years for orbit decay and atmospheric reentry in guideline 6-1a because retrieval leaves the space system in high value regions of space, whereas transfer to an orbit with reduced lifetime lowers the perigee of the final mission orbit and reduces the fraction of time spent in high value regions of space.

Using guideline 6-2a, the region of space above GEO altitude can be used as a disposal region with little concern for debris buildup because of the low relative velocities, large regions of available space, and relatively low traffic rates in this area. In the near future, the 300 km altitude separation will be sufficient to isolate the disposal region from GEO if steps are taken to remove on-board energy sources after completion of the postmission disposal maneuver (guideline 4-1). However, depending on the level of traffic to GEO and on the characteristic sizes of future GEO satellites, this separation distance may need to be increased in the future. If measures are not taken to prevent explosive structural failure after disposal of GEO systems, a separation distance of ~2000 km will be required to isolate the disposal region from GEO.

The four burns prescribed in guideline 6-2a take into account fuel gauging limitations which are particularly serious at the end of missions. Given that there will be significant uncertainty as to the amount of fuel remaining at the end of mission, there will also be some uncertainty as to whether there is enough fuel to complete the disposal maneuvers. If there is not enough fuel to complete the maneuvers, a plan using four smaller burns to maneuver to the disposal orbit will leave the spacecraft or upper stage farther from GEO altitude than would a plan using two burns. In planning the postmission disposal, uncertainties in fuel gauging should be considered in the assessment of reliability.

In guideline 6-3, a program with a near-circular, 12-hour orbit will need to maneuver to a disposal orbit above or below the semisynchronous region. The separation from semisynchronous circular orbit is at least 300 km, so lunar and solar perturbations will not cause these objects to interfere with spacecraft operating in 12-hour orbits.

To satisfy guideline 6-4, systems should be removed from useful regions of space with a high probability of success. A reliable propulsion system will have a probability of failure on the order of 0.01. Failure of the propulsion system should be the primary source of failure to perform successful postmission disposal.

The amount of time a structure will remain in orbit depends on its final orbit and on the area-to-mass ratio, as discussed in chapter 3. The evaluation in this section follows that developed for guideline 3-1.

The steps in the evaluation are described below.

Spacecraft using atmospheric drag and reentry for postmission disposal need to be evaluated for survival of structural fragments to the ground. Guidelines for this evaluation are presented in chapter 7 of this standard.

All other disposal options result in the space system being left in long-lifetime orbits that will not interfere with future space operations. A plan for performing the postmission disposal maneuvers should be included in the assessment.

The debris assessment should consider two areas: (1) design or component failure which leads to loss of control during the mission, and (2) failure of the postmission disposal system, including insufficient fuel to complete the disposal operation. Conventional failure modes and effects analysis or equivalent analysis can be used to assess failures which lead to loss of control during mission operations and postmission disposal.

Note: The probability of damage from collision with orbital debris or meteoroids leading to loss of control is analyzed under guideline 5-2.

For a program or project that elects to limit orbit lifetime using atmospheric drag and reentry, there are several options for reducing orbit lifetime:

4. To increase the probability that the postmission disposal maneuver will be successful, the program may want to consider incorporating redundancy into the postmission disposal system.

Programs or projects that use atmospheric reentry to limit the orbit lifetime of their systems in conformance to guideline 6-1 present a potential risk to the Earth's population. This chapter presents the guideline that defines the maximum amount of debris that can survive reentry if the reentry is uncontrolled. Uncontrolled reentry is defined as reentry in which the ground footprint location cannot be determined with sufficient accuracy to guarantee missing landmasses.

The guideline for uncontrolled reentry provides an upper limit of 8 m² on the total casualty area of debris that impacts the Earth. An upper limit of 8 m² is derived by assuming an average risk of human casualty of 0.0001 per reentry event. However, the risk of a reentry event causing any casualties is actually lower since no correction has been made for the fact that people are usually protected inside buildings or vehicles and will therefore be shielded from reentering debris. To date, no casualties have been attributed to reentering man-made space structures.

The measure of risk from reentering debris is the debris casualty area. For a piece of debris that survives atmospheric reentry, the debris casualty area is the debris cross-sectional area plus a factor for the cross-section of a standing individual. The total debris casualty area for a reentry event is the sum of the debris casualty areas for all debris pieces surviving atmospheric reentry. Equation 7-1 is used to calculate the total debris casualty area.

An approach for estimating the total debris casualty area of a space structure from a decaying orbit is summarized in figure 7-1. The following procedure is used to determine if a reentering space structure exceeds the total debris casualty area limits. The parent body is the structure as it exists in orbit.

(Note: This procedure has been automated in the debris assessment software)

Figure 7-1. Flow chart for calculating total debris casualty area, D_A, for uncontrolled reentry.

[J/kg]

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