Re-entry Analysis

Drag augmented semi-controlled re-entry: Proof of concept

Uncontrolled re-entries of satellites and rocket bodies can pose an undue risk to humans on Earth due to surviving fragments. With current technologies, uncontrolled re-entries can be tolerated for spacecraft with a dry mass up to one ton. With proper design measures (“Design-for-Demise”), this limit can probably be extended to two tons. Above two tons, controlled re-entries, targeting for uninhabited areas (e.g. the South Pacific Ocean), have to be used in order to control the on-ground risk. Controlled re-entries can cause significant additional costs for a satellite project in the order of several million euros (development, manufacturing, launch, and operational costs).

One of the investigated self-removal technologies of the TeSeR project makes use of a drag sail deployment at quite low altitudes (<200 km). Such drag sails should be large enough to generate sufficient drag force for a semi-controlled re-entry. In contrast to the fully controlled re-entry, where the targeted impact zone has a length in the order of just a few thousands of kilometers, a semi-controlled re-entry is targeting for an impact zone with a length up to a couple orbits. Risk reduction is achieved by selecting impact orbit arcs with minimum population density.

In order to confirm the general feasibility of drag-assisted removal concepts resulting in a semi-controlled re-entry, the analysis was focused on the following three main topics:

  1. Residual lifetime estimations
  2. Identification of minimum populated ground-track arcs as potential targets for a semi-controlled re-entry
  3. Determination of thermal and mechanical loads on the drag sail after deployment

The length of the impact orbit arc depends on the residual lifetime until re-entry. Uncertainties for the atmospheric density are causing uncertainties for the residual lifetime in the order of up to ±20%. For example, 20 hours of residual lifetime corresponds to a re-entry window size of ±96 minutes or ±1 orbit.

The goal was to find solutions for one- or two-ton class satellites. Drag sails with an area of 100-200 m² should be feasible for such large satellites. Thus, the area-to-mass ratio of interest is about 0.1 m²/kg. The residual lifetime for such a system (until 90 km altitude) estimated with a numerical propagator is in the order of 50 minutes for drag sail deployment at 150 km altitude. In order to be safe, we decided to increase the residual lifetime uncertainty to ±50%. In this case, the achievable re-entry arc length is 100 minutes (now also including the time from 90 km down to ground).

Based on population density databases, a tool was developed to find the location of the minimum populated arc for a given length and orbit inclination. Figure 1 shows the result for a polar orbit. Compared to an uncontrolled re-entry, this semi-controlled re-entry would reduce the on-ground risk to only 0.4%.

However, it is quite unlikely that the drag sail will remain intact until an altitude of 90 km. Additional re-entry simulations showed that drag sails based on Nylon will probably fail already at about 145 km altitude due to exceeding of the maximum temperature limit of the sail material. Figure 2 shows the orbit evolution of the sample mission (one ton satellite, 100 m² drag sail) before and after the sail failure event.

Although the sail fails quite early, aerodynamic capture and residual lifetime reduction is achieved, especially if the sail is deployed at 150 km altitude. Nevertheless, the total time until ground impact is only slightly higher than previously estimated: 116 mins from deployment at 150 km altitude. This extends the minimum populated arc, but does not affect the achievable risk reduction. The needed extra arc length is added over uninhabited areas.




Removal Strategies

Removal strategies for the post-mission-disposal (PMD) of spacecraft vary.

The TeSeR project investigates three removal concepts:

  1. Uncontrolled
  2. Semi-controlled
  3. Controlled


User and Mission Requirements

User and Mission Requirements

The user and mission requirements are derived from a host of national and international standards and regulations.


While the system of international space treaties (e.g. the Outer Space Treaty of 1967) provide rather rough guidelines on space debris, more recently the UN Committee on the Peaceful Uses of Outer Space – particularly the Working Group on the Long-term Sustainability of Outer Space Activities – has drawn up a much more comprehensive and concrete set of guidelines.

Based on the guidelines of the Inter-Agency Space Debris Coordination Committee (IADC) two branches of recommendations and requirements were established. A) the European Code of Conduct for Space Debris Mitigation and B) ISO 24113 and ECSS-U-AS-10C.


The European Code of Conduct for Space Debris Mitigation is a voluntary agreement between European national space agencies and ESA. It shall ensure that no new space debris is generated by future missions and it is mandatory for ESA projects with the requirements on space debris mitigation for ESA projects.

Within Europe, ECSS-U-AS-10C is now the primary source of debris mitigation requirements. ESA updated its Space Debris Mitigation Policy document in 2014 to adopt ECSS-U-AS-10C as its standard for space debris mitigation.


All these requirements influence national standards and reports which have again an influence on the requirements itself.

Debris Guidelines

The TeSeR post-mission-disposal (PMD) module shall be compliant to ECSS and ISO.

Based on the user requirements a list of high-level mission requirements is derived.

TeSeR Module Specification

Within the TeSeR project, the work objectives of Bundeswehr University Munich were the development of concepts and the definition of a functional architecture of the post-mission-disposal (PMD) module.

It is critical to discern when a spacecraft is not operating nominally. Therefore, as a first step of the analysis, it was investigated how this can be detected. A number of symptoms indicating a troubled spacecraft are identified (see figure 1). The majority of the symptoms can only be checked through checks of the telemetry of the host spacecraft. This is done through the ground station. In addition, certain physical parameters can indicate a non-nominal state of the host spacecraft: excessive rotational rates, temperatures exceeding acceptable limits, and lack of power. In order to detect these symptoms from the PMD module, a suitable set of sensors is identified. As an added bonus, this allows the PMD module to establish the position and attitude of the host spacecraft if needed.

1 - detection_V01mk

Figure 1: Spacecraft status detection

The operational and autonomy concept is aimed at “triggering” the removal from orbit and the passivation of the host spacecraft (i.e. safely disabling) at the end of its operational lifetime. It was found that future spacecraft using the PMD module technology will be required to be designed to be self-passivated. The PMD module itself will also have to be passivated having performed the removal operations unless in case of a direct, controlled atmospheric re-entry. The operational autonomy concept foresees – for the foreseeable future – a human operator “in the loop”. However, in the case of satellite mega-constellations, autonomous removal presents a viable business case as it can significantly reduce unnecessary cost. The autonomy concept features three levels of increasing on-board autonomy capabilities (see figure 2).


2 - Autonomy levels

Figure 2: Illustration of the three progressively more advanced autonomy levels defined by Bundeswehr University Munich for removal from orbit

The removal subsystems are placed on the PMD module platform. The PMD module platform is placed on the host spacecraft. The system design comprises of the concept of operations for the PMD module (see figure 3), the functional design (“What the system should do.”), and the performance design (“How well the system should do it”). The established requirements for the design will form the basis for future, most likely modular, more detailed designs. Regarding the level of self-sufficiency of the PMD module, the analysis showed that the optimal route is to not be reliant on the host spacecraft.


Figure 3: The top-level operational concept for removal from orbit as defined by Bundeswehr University Munich

4 - Functional block diagramFunctional block diagram hi res


Figure 4: A functional block diagram of a version of the PMD module, its subsystems, the connections between them, and the two interfaces to the host spacecraft and the removal subsystem (RS) respectively

For the interfaces between the host spacecraft and the PMD platform as well as the PMD platform and the removal subsystem, the approach follows the idea to standardize as much as possible (see figure 5). Regarding the physical/mechanical, thermal, electrical, and data interface to the spacecraft, the aim was to achieve minimal reliance on the resources of the spacecraft. As the PMD module platform will “trigger” the removal subsystems via Standard Interface #2, this “fuller” interface encompasses a larger set of options.


PMD module concept

Figure 5: The concept of the PMD module: the PMD module platform is attached to the host spacecraft via standardized interface #1. Each removal subsystem can be attached to the PMD platform via standardized interface #2