Case Study Planetary Protection Category

Case Study

Planetary Protection Category III

Presented by Dr. Gerhard Kminek, COSPAR Table of Content

 Planetary protection category III description  Case study for planetary protection category III  Requirements for case study  Implementation of requirements for case study  Things to remember Category III description

Flyby (i.e. gravity assist) and orbiter missions to a target body of chemical evolution and/or origin of life interest and for which scientific opinion provides a significant2 chance of contamination which could compromise future investigations

2Implies the presence of environments where terrestrial organisms could survive and replicate, and some likelihood of transfer to those places by a plausible mechanism

Applicability: Mars, Europa, Enceladus

Credit: ESA/Mars Express Credit: NASA/JPL/Galileo Credit: NASA/JPL/Cassini

However, if an orbiter mission is looking for life, the mission will have to meet requirements for a life detection mission (i.e. avoid compromising the life detection measurement)

Credit: NASA/JPL/Cassini Case study

ExoMars Trace Gas Orbiter (TGO)

→ Target body: Mars

→ Propulsion: Chemical

→ Transfer: Deterministic Deep Space Manoeuvre

Credit: ESA/ExoMars (DSM) with several 100 m/s and stochastic

Trajectory Correction Manoeuvres (TCMs) with

several m/s

→ Orbit acquisition: Mars Orbit Insertion (MOI)

manoeuvre with several 100 m/s and aerobraking

→ Final orbit: 400x400 km, 373:30 repeat pattern Requirements for case study Launcher upper stage The probability of impact on Mars by any element not assembled and maintained in ISO level 8 conditions shall be ≤ 1x10-4 for the first 50 years after launch

Note: This requirement also applies if a launch service is provided to another customer, e.g., launch of the Emirates Mars Mission (EMM) on H-IIA

Spacecraft One of the following conditions shall be met: → The probability of impact on Mars by any part of a spacecraft assembled and maintained in ISO level 8 cleanrooms, or better, is ≤ 1x10-2 for the first 20 years after launch, and ≤ 5x10-2 for the time period from 20 to 50 years after launch (e.g., Mars Express, TGO)

→ The total bioburden of the spacecraft, including surface, mated, and encapsulated bioburden, is ≤ 5x105 bacterial spores (e.g., MRO, Maven) Implementation of requirements

Launcher upper stage The probability of impact on Mars by any element not assembled and maintained in ISO level 8 conditions shall be ≤ 1x10-4 for the first 50 years after launch*

*Relevant requirement needs to be reflected in Launcher Interface Requirements Trajectory analysis based on Monte Carlo method to achieve a one-sided 99% level-of- confidence (Wilson interval) → Trajectory analysis covers all reference trajectories for the launch window (i.e. >1) → Number of Monte Carlo runs depends on the detected number of impacts (iterative)

Analysis includes: → Gravity potential of Earth and Mars and 3rd body perturbation by Sun, Moon, Jupiter and Saturn → Solar radiation pressure (SRP) with un-controlled attitude → Propellant blow-down as directed contribution, outgassing as spherical contribution → In case there is a manoeuvre of the upper stage after the release of the spacecraft (e.g., for Breeze-M and Centaur), the reliability of this manoeuvre (from flight records) has to be part of the overall analysis

In case impact probability is too high  increase launch bias away from Mars (effect on delta-v budget for spacecraft) Implementation of requirements

→ The total bioburden of the spacecraft, including surface, mated, and encapsulated bioburden, is ≤ 5x105 bacterial spores Implementation of requirements

Credit: ESA/ExoMars 1. Analyse the stability of the final science orbit  stable for >50 years √ • Numerical propagation of orbit with atmospheric variation (driven by solar cycle); evaluate right ballistic parameter and proper parameters from the atmospheric model 2. Analyse the impact probability of the spacecraft before the DSM  no impact >50 years √ • Monte Carlo method to achieve a one-sided 99% level-of-confidence (Wilson interval)

• Necessary input is the launcher dispersion matrix (injection conditions) • Gravity potential of Earth • 3rd body perturbation by Sun, Moon, Jupiter and Saturn Credit: ESA/ExoMars • Solar radiation pressure (SRP) with controlled and un-controlled attitude 3. Analyse the impact probability between the DSM and reaching the final science orbit • Assume trajectory impact probability of ‘1’ after DSM (conservative) • Reliability of the flight hardware necessary to control the spacecraft and reliability of operation • Atmospheric variation for Mars aerobraking phase; ignore chance for recovery (conservative) • Micrometeoroid impact and effect analysis (details next) Credit: ESA/ExoMars

-2 Probability of impact: PHW failure + POP failure + Pmeteoroid kill ≤1x10 Implementation of requirements

1. Micrometeoroid model definition → Selection of micrometeoroid flux model (e.g., Grün), velocity distribution (e.g., 20 km/s), micrometeoroid density (e.g., 2.5 g/cm3), and average impact angle (e.g., 45°)

2. Analysis of consequences → Selection of critical units necessary to control spacecraft (see reliability analysis)

→ Assess protection based on presence of MLI, panels, honeycomb panels, etc. in terms of equivalent thickness; take into account view factors redit: ESA/ExoMars C → Assess protection based on distances between different elements MLI CFRP SLI

to select use of proper ballistic limit equation (BLE from IADC) → Assess the failure modes of critical hardware Rear Wall → Typical problematic hardware e.g., tanks, star trackers, propulsion Bumpers lines, UHF RFDN waveguides Credit: ESA/ExoMars Implementation of requirements  The approach used for the TGO is conservative in many ways but also easier to evaluate  This approach can be used for orbiter missions and for cruise stages delivering a lander

 In case this approach is not sufficient to demonstrate compliance with the probability of impact requirements, other approaches could be used (but do not guarantee compliance) → Replacing the fixed micrometeoroid velocity with a velocity distribution (e.g., Taylor) → Replacing the trajectory impact probability value of 1 during PI: prob. of impact during this phase th part of the cruise phase with a value based on a trajectory PI = Σipiqi+1 pi: prob. of impact due to the i manoeuvre qi+1: prob. that the next manoeuvre to remove analysis and allowing recovery manoeuvres the impact fails → Allowing recovery manoeuvres during aerobraking

PI = Σi(orbit)[pi(non-recoverable)+pi(recoverable)qi]+Σj(manoeuvre)[pj(non-recoverable)+pj(recoverable)qj]+P(safe mode)+Q(t1, t2)

Σi(orbit)[pi(non-recoverable) + pi(recoverable)qi]: prob. of impact of a healthy spacecraft due to random variations of the atmosphere Σj(manoeuvre)[pj(non-recoverable) + pj(recoverable)qj]: prob. of impact of a health spacecraft due to tracking, manoeuvre & operational issues P(safe mode): prob. to enter a safe mode Q(t1, t2): prob. of impact due to hardware reliability and micrometeoroids Things to remember

 Probability of impact requirements can have an effect on the qualification of hardware (e.g., solar arrays for aerobraking), the trajectory design, the delta-v budget (re-targeting), and spacecraft design (e.g., location of tanks, additional micrometeoroid protection)  To accommodate these effects, have a first analysis ready for the PDR  This first analysis should not be too simplistic – otherwise late changes in the spacecraft design or operation might become necessary  There is a trade-off in the aerobraking design between more gentle and longer aerobraking (negative for micrometeoroid effects and reliability) and more aggressive and shorter aerobraking (negative for hardware qualification and operation)  Ensure good interface with launcher system for upper stage impact analysis  All activities necessary to perform a probability of impact analysis are interdisciplinary and require the interactions between different engineering disciplines!