Describing the Mission
The simplest way to define the goals of Phootprint (the Phobos landing and sample retrieving mission), would in layman terms be,
To safely reach Phobos, land a probe to collect samples, send the samples back to Earth for study without contamination, leakage or threats to either the samples or Earth.
Describing the mission objectives in detail would be: –
• Return approximately 100g of loose material from the surface of Phobos.
• Access at least 50% of the Phobos surface for the sampling operations Landing site selected by Science team during mission post extensive global and local mapping campaign (grooves of particular interest and drove 100m major axis landing accuracy).
• Maximize science return on the returned sample Strict requirements on surface contamination.
• Goal of 800g load on sample.
• Static landing (no touch-and-go) Allows for more precise selection of sampling point.
• Target of 2024 launch with 2025/2026 backup Ariane 5 ECA baselined.
• Earth swing-by provides greater launcher margin.
In order to identify Our goals in detail, it is common that a series of questions is asked, questions that help us clearly understand our goals such as:
• What is our major scientific goal in this mission?
• What are our engineering prospects that are required and that we need to fulfill?
• Why are we collecting samples from Phobos?
• Why Phobos?
• What would be our future aims and studies using the results of this mission (in other words: what is the Phootprint of this mission)?
• What is Phobos and how do we describe and tackle it?
• And eventually our major Goal in this study, how do we go about performing the trade-off analysis for the mission and the space seal?
These goals are internationally set by ESA (the European Space Agency) and the missions major contributing countries, Namely: England, France and Germany.
We will discuss briefly for now the goals shortly for the mission. Asking ourselves questions such as, what is our major scientific goal in this mission? What are our engineering prospects that are required and that we need to fulfill? Why are we collecting samples from Phobos? Why Phobos? What would be our future aims and studies using the results of this mission (in other words: what is the Phootprint of this mission)? What is Phobos and how do we describe and tackle it? And eventually our major Goal in this study, how do we go about performing the trad-off analysis for the mission and the space seal?
The mission objective and preliminary procedures in detail: –
Phootprint (or Phobos Sample Return) mission will return samples from Mars satellite Phobos. These samples will be transported in a container encapsulated in an Earth Re-entry Capsule (ERC) that will perform a hard landing on Earth. The ERC will be released from the Earth Return Vehicle (ERV) and re-enter earth’s atmosphere. The re-entry stage will subject the ERC to demanding thermo-mechanical loads. Through all the process the ERC will be responsible for ensuring the sample container is below established levels of temperature and deceleration.
The samples shall not only try to cover ground material, but as much of the gaseous or liquid material possible. Hence, along with this comes the necessity to align the load factors appropriately in such a way that they are layered in such a way that none of the samples is bothered by the loads.
The Landing procedure, will not be a simple one since not much can be predicted or described about the surface of Phobos. Hence, a much complex landing procedure will be implied. A so called “Static landing” described as a no-touch and go landing I.e. the probe won’t fully come in contact with the surface of the outer layers of Phobos making it much harder to collect the samples as in the have to be from very specific entry and exit (or collection points and zones) from the probe, requiring a very specific design and engineering prospects.
Required characteristics of the mission and sampling probe: –
Engineering prospects of the sampling probe and seal: –
Hard landing on Earth of an ERC without the use of parachutes or propulsion systems is a very demanding thermomechanical situation for which the selected materials should be thermomechanical compliant with the mechanical and thermal efforts developed during all the re-entry phase.
For the re-entry phase a TPS material is introduced in the ERC to limit the heat conduction to the cold structure from the ablative layer. Also, it controls the temperature at the interface with the inner compartment (bio container and beacons). In the case of passive landing it is needed an interface material that dissipate/absorb the kinetic energy controlling, simultaneously, the container’s deceleration and maintaining the thermal conduction as low as possible. One way to control container´s deceleration is the gradual compression of the interface material what could be achieved through the gradual compression of a crushable foam.
A crushable thermal protection system, cTPS, shall perform simultaneously structural and thermal functions, according to the space application. The former requires the capability of: a) absorption of impact energy during landing; b) sustaining dynamic pressure induced loads during re-entry; and c) other mechanical requirements. The thermal function must provide the capability of: a) sustaining the aerothermodynamics heat loads during re-entry; b) keeping the sample canister (or bio-container) at a specified temperature; and, c) other thermal requirements. The crushable TPS system together should be able to protect the bio-container from high decelerations peaks and extreme temperatures during ERC’s re-entry, impact and post-impact phases.
The activities are developed under the ESA project: Design of a Crushable TPS for the ERC (TECMSS/2011/226/ln/IN) with the principal aim to investigate and design a multifunctional cTPS structure that act as a heat shield for planetary re-entry and brings an adequate impact response during hard landing. Two different ERC’s were considered for development. The first ERC is named MSR with an external diameter equal to approximately 1.5m and maximum weight varying in the range 120, 150kg. For this geometry two types of thermal protection systems materials (varying material’s density and thermal properties) named LD-TPS and HD-TPS were considered. The Phootprint ERC is a smaller version of the MSR’s ERC with 0.75m of external diameter and 30kg in weight. Below are summarized the main objectives for the design and production sequence:
• identify materials and mixed solutions for thermal and structural protection;
• material characterization for design concept and analysis;
• set up a conceptual solution for cTPS solution;
• design and built breadboard models for concept demonstration;
• reach to the TRL 3 level in the concept design and experimental demonstration;
• perform experimental tests to verify the proposed cTPS systems.
Phobos is the innermost and larger of the two natural satellites of Mars the other being Deimos. Both satellites were discovered in 1877 by American astronomer Asaph Hall. It is a small, irregularly shaped object with a mean radius of 11 km (7 mi) and is seven times as massive as the outer satellite, Deimos. It is named after the Greek god Phobos, a son of Ares (Mars) and Aphrodite (Venus), and the personification of horror (phobia).
Phobos orbits 6,000 km (3,700 mi) from the Martian surface, closer to its primary body than any other known planetary satellite. It is indeed so close that it orbits Mars much faster than Mars rotates and completes an orbit in just 7 hours and 39 minutes. As a result, from the surface of Mars it appears to rise in the west, move across the sky in 4 hours and 15 minutes or less, and set in the east, twice each Martian day.
Phobos is one of the least reflective bodies in the Solar System, with an albedo of just 0.071. Surface temperatures range from about ?4 °C (25 °F) on the sunlit side to ?112 °C (?170 °F) on the shadowed side. The defining surface feature is the large impact crater, Stickney, which takes up a substantial proportion of the satellite’s surface.
Images and models indicate that Phobos may be a rubble pile held together by a thin crust, and that it is being torn apart by tidal interactions. Phobos gets closer to Mars by about 2 meters every one hundred years, and it is predicted that within 30 to 50 million years it will either collide with the planet or break up into a planetary ring.
Phobos has dimensions of 27 km × 22 km × 18 km and retains too little mass to be rounded under its own gravity. Phobos does not have an atmosphere due to its low mass and low gravity. It is one of the least reflective bodies in the Solar System, with an albedo of about 0.071. Spectroscopically it appears to be similar to the D-type asteroids and is apparently of composition similar to carbonaceous chondrite material. Phobos’s density is too low to be solid rock, and it is known to have significant porosity. These results led to the suggestion that Phobos might contain a substantial reservoir of ice. Spectral observations indicate that the surface regolith layer lacks hydration, but ice below the regolith is not ruled out.
Phobos is heavily cratered. The most prominent of these is the crater, Stickney, a large impact crater some 9 km (5.6 mi) in diameter, taking up a substantial proportion of the satellite’s surface area. As with Mimas’s crater Herschel, the impact that created Stickney must have nearly shattered Phobos.
Many grooves and streaks also cover the oddly shaped surface. The grooves are typically less than 30 meters (98 ft.) deep, 100 to 200 meters (330 to 660 ft.) wide, and up to 20 kilometers (12 mi) in length and were originally assumed to have been the result of the same impact that created Stickney. Analysis of results from the Mars Express spacecraft, however, revealed that the grooves are not in fact radial to Stickney, but are centered on the leading apex of Phobos in its orbit (which is not far from Stickney). Researchers suspect that they have been excavated by material ejected into space by impacts on the surface of Mars. The grooves thus formed as crater chains, and all of them fade away as the trailing apex of Phobos is approached. They have been grouped into 12 or more families of varying age, presumably representing at least 12 Martian impact events. Faint dust rings produced by Phobos and Deimos have long been predicted but attempts to observe these rings have, to date, failed. Recent images from Mars Global Surveyor indicate that Phobos is covered with a layer of fine-grained regolith at least 100 meters thick; it is hypothesized to have been created by impacts from other bodies, but it is not known how the material stuck to an object with almost no gravity. A 68 kg (150 lbs.) person standing on the surface of Phobos would weigh the equivalent to about 60 g (2 oz.) on Earth.
Temperature variation on the surface of Phobos is as shown above: –
Orbital characteristics: –
The orbital motion of Phobos has been intensively studied, making it “the best studied natural satellite in the Solar System” in terms of orbits completed Its close orbit around Mars produces some unusual effects. With an altitude of 5,989 km (3,721 mi), Phobos orbits Mars below the synchronous orbit radius, meaning that it moves around Mars faster than Mars itself rotates. Therefore, from the point of view of an observer on the surface of Mars, it rises in the west, moves comparatively rapidly across the sky (in 4 h 15 min or less) and sets in the east, approximately twice each Martian day (every 11 h 6 min). Because it is close to the surface and in an equatorial orbit, it cannot be seen above the horizon from latitudes greater than 70.4°. Its orbit is so low that its angular diameter, as seen by an observer on Mars, varies visibly with its position in the sky. Seen at the horizon, Phobos is about 0.14° wide; at zenith it is 0.20°, one-third as wide as the full Satellite as seen from Earth. By comparison, the Sun has an apparent size of about 0.35° in the Martian sky. Phobos’s phases, since they can be observed from Mars, take 0.3191 days (Phobos’s synodic period) to run their course, a mere 13 seconds longer than Phobos’ sidereal period. As seen from Phobos, Mars would appear 6,400 times larger and 2,500 times brighter than the full Satellite appears from Earth, taking up a quarter of the width of a celestial hemisphere. The Mars–Phobos Lagrange L1 is 2.5 kilometers (1.6 mi) above Stickney, which is unusually close to the surface.
The origin of the Martian satellites is still controversial. Phobos and Deimos both have much in common with carbonaceous C-type asteroids, with spectra, albedo, and density very similar to those of C- or D-type asteroids. Based on their similarity, one hypothesis is that both satellites may be captured main-belt asteroids. Both satellites have very circular orbits which lie almost exactly in Mars’s equatorial plane, and hence a capture origin requires a mechanism for circularizing the initially highly eccentric orbit, and adjusting its inclination into the equatorial plane, most probably by a combination of atmospheric drag and tidal forces, although it is not clear that sufficient time is available for this to occur for Deimos. Capture also requires dissipation of energy. The current Martian atmosphere is too thin to capture a Phobos-sized object by atmospheric braking. Geoffrey Landis has pointed out that the capture could have occurred if the original body was a binary asteroid that separated under tidal forces.
Phobos could be a second-generation Solar System object that coalesced in orbit after Mars formed, rather than forming concurrently out of the same birth cloud as Mars.
Another hypothesis is that Mars was once surrounded by many Phobos- and Deimos-sized bodies, perhaps ejected into orbit around it by a collision with a large planetesimal. The high porosity of the interior of Phobos (based on the density of 1.88 g/cm3, voids are estimated to comprise 25 to 35 percent of Phobos’ volume) is inconsistent with an asteroid origin. Observations of Phobos in the thermal infrared suggest a composition containing mainly phyllosilicates, which are well known from the surface of Mars. The spectra are distinct from those of all classes of chondrite meteorites, again pointing away from an asteroid origin. Both sets of findings support an origin of Phobos from material ejected by an impact on Mars that reaccreted in Martian orbit, similar to the prevailing theory for the origin of Earth’s satellite.
Phootprint Mission Overview Schematics and Imagery.
Described within the next image, is the real time launch statistics and how the trajectory of the probe would progress in real time described within the axis and time with dates in earth days.
Such estimations are done using calculatuons and simulations to estimate certain and various aspects of the mission in order to ensure the smooth sailing and running of the mission at various instants in time.
One of the major steps to prepare the probe would be to design and analyze the systematic and physical properties of it in order to ensure that the required physical objectives of the probe are fulfilled efficiently. Along with such, we have the next image, which is the first simulation of the features required and the first 3D model defined for the Phootprint probe.
After the 3D model of the entire structure has been described, we move on to analyze the different features of it for different layers and parts of the entire probe. Starting off with dissecting the outer hull of the structure and defining its inner parts in separate and different parts. This helps us understand more the physical characteristics of the probe and whether it will withstand the required objectives, allowing us to reform our structure, redesign or overhaul the entire blueprint if required.
Here, we will describe as much as possible the overview of the trajectory without real time requirements or analysis to define the mission in the simplest stages using entries, exits, landings etc.
The mission starts with an ignition on the surface of Earth. If ignition successful, Lift-off is obtained. Implying that the rocket launched is no longer under the influence of gravity from Earth. After exiting the exosphere, the probe will be in free space vacuum, after escaping completely the gravitational influence of Earth, we are now obliged to use high energy propulsion systems to maintain the trajectory. After reaching proximity of mars, we will attempt the landing using the gravitational acceleration and force of mars to enter and orbit which will allows to land safely and without damage on phobos without collision. After the sampling is complete, we will discard all weighting and non-essential exit and re-entry equipment i.e. we will let go off all non-required equipment. We will use the thrusters to exit the gravitational influence of phobos and use mars to enter the earth transfer orbit. We will once again use high propulsion systems to reach proximity into earth and use the force of the re-entry of the atmosphere and gravitational pull of earth to collide on earths surface.
Designing and testing the sealed capsule.
In order to choose the best possible material and design, firstly we shall set onwards the certain requirements that the sealed capsule is required to have in order to get the best results.
Introducing our Requirements: –
The requirements of the design will be described, and set based on previous missions and expeditions that involve both probing and sample collection. Addressed previously in the “Evaluation of Sealing Systems for a Phobos Sample Return Mission” paper addressed by the comoti research center in Bucharest, Romania targeting to design, manufacture and test a leak tight container from now on addressed as “the vault” that shall accommodate the sample canister altogether with the 100g sampled regolith material.
In the frame of the previous projects the first task was to analyze the system requirements and the main two proposed sampling concepts for the Phobos Sample Return Mission. Based on the reference documents provided by ESA, a critical review of the system requirements was conducted using several factors like technological implications, material impact, functional implications, design points and others in order to better understand and assimilate them, and also two new requirements were proposed. The two sampling concepts provided by Airbus and Thales were analyzed and improvement areas regarding the sampling process and the sealing system were identified and proposed for each of them. The assessment of the two sampling concepts consisted of the concept description, concept analysis based on the requirements, emphasizing the strong points and weaknesses for each concept and identifying areas of improvement both for the overall concepts and for the weaknesses.
Based on the information from the first task, the second task was to analyze several types of sealing systems with respect to the Phobos Sample Return Mission, to identify the critical technologies that require pre-development and to conduct a trade-off analysis of the selected technologies, analysis that will provide a baseline for the design phase. First of all, the sealing technologies were classified in two main categories: fuse sealing methods and pre-load sealing methods. For pre-load sealing methods were considered and analyzed O-rings, gaskets, knife edge sealing systems and shape memory materials, while for the fuse sealing methods were considered and analyzed crimp, vibration/ultrasonic welding, soldering, brazing and adhesive sealants. The main difference between the two categories is that the pre-load sealing method do not require a back-up aperture for opening on Earth. Each sealing method was analyzed according an imposed format consisting of the following criteria: introduction and general presentation, advantages and disadvantages, relevant values for the mission, other uses of the method in space applications and finally the conclusion.
Real physical technical requirements: –
Imposed and dictated by the outcome of previous studies, experiments and written handbooks of rules and regulations, the following 27 technical requirements were proposed: –
1. The sample container vault shall be cylindrical in shape
2. The sample container design shall be taken from the AVS study
3. The design of the vault shall avoid, as far as possible, contamination of the sample or sample container during the opening procedure once the ERC is returned to the receiving facility on Earth.
4. The design of the vault shall be such as to minimize the control requirements on the robotic arm during insertion of the sample container into the vault.
5. The sample container vault shall be sealed using a maximum force of 40 N and a maximum torque of TBC Nm.
6. The sealing technologies to be considered shall be at the minimum (TBC): O-ring, metallic Gasket, Knife edge.
7. A redundant and fail-safe containment sealing system shall be envisaged for the vault.
8. The vault (excluding sample container) and supporting structures (i.e. brackets to attach to the ERC structure) shall have a maximum mass of 2.0kg including 20% margin.
9. The vault sealing system shall preserve its functionality during all its life.
10. The Vault shall be designed to be installed into the ERC.
11. Testing of the closure of the sealing system shall be performed under thermal and vacuum conditions to verify the tightness of the seals.
12. The sealing shall be verified under relevant environmental conditions for the mission profile between closure and just prior to Earth impact (duration to be accelerated).
13. Shock testing of full-scale breadboards shall be included in the test program.
14. Tests shall be performed to demonstrate robustness of the seal under representative but varying levels of contamination using Phobos analogue materials.
15. Testing of any launch locks under launch loads specified in REQ-17 to REQ-21 below shall be performed.
16. The vault shall withstand all the mechanical environmental loads encountered during its lifetime (i.e. launch, re-entry and landing).
17. The 1st eigenfrequency of the vault shall be ; 180 Hz (TBC), in hard mounted conditions.
18. The vault shall withstand the following design / qualification level quasi-static accelerations:
a. X-axis 25 g
b. Y-axis 25 g
c. Z-axis 25 g
19. The vault shall withstand the following qualification level sine vibration environments:
a. X, Y and Z-axis
b. 5 Hz 1 g
c. 5 – 30 Hz linear slope
d. 30 – 100 Hz 25 g
20. Sine vibration environments applied independently in X, Y and Z directions. Sine sweep up and down in each axis between 5 and 100Hz at 2 oct/min.
21. The vault shall withstand the following qualification level random vibration environments:
22. X, Y and Z –axis (11.1 gRMS)
a. 20 – 40 Hz +6 dB/oct
23. – 450 Hz 0.16 g2/Hz
a. s450 – 2000 Hz -6 dB/oct
24. Note: Random vibration environments applied independently in X, Y and Z directions, with a duration of 2.5 min per axis
25. The vault shall withstand the following enveloping qualification SRS shock environment (Q=10).
a. 100 Hz 25 g
b. 2000 Hz 1500 g
c. 10000 Hz 1500 g
26. The vault shall withstand the quasi-static loads during Earth re-entry within a limit of TBC g.
27. The vault shall withstand a shock load of 2000g for 10 ms without failure.
28. The vault and sealing systems shall be compatible with cleaning processes that reduce organic and particulate contamination to a level of TBC.
29. The vault shall withstand the thermal environment encountered during the lifetime of the mission (operating temperatures and non-operating temperatures).
30. The vault shall be vacuum compatible (TML100g.
4. SR 4 – Before closure of the ERC the sampling system shall be capable of at least one more sampling attempt in case of failure of the first sampling attempt. Potentially this could lead to a more severe particle contamination of the seals. Update: The sealing and locking system shall work only once (it will not be opened again; the second sampling attempt takes place before closing the ERC).
5. SR 5 – The sampled material consists of a grain size distribution between 10 years.
11. SR 11 – The landing can occur up to 20 degrees off axis.
Based on the identified test plan requirements a set of guidelines for each individual test is presented below.
Leakage test: – SR1
– No particle or droplet of a fluid ; 1?m shall escape or enter the sample canister;
– The sealing system must be active in both directions;
– The system must be leak tight for all expected environmental conditions and all types of tests performed;
– Information about the equipment used in order to prove that the SR1 is met will be provided;
– During and after the manufacturing, the components shall be subjected to non-destructive tests.
Shock test: – SR2, SR3, SR11
– The seal shall withstand a shock of 2000 g for 10 ms without failure;
– The amount of 100g of regolith collected must be taken into account because at 2000 g the inertial forces are considerable high;
– It shall be taken into account the fact that the landing can occur up to 20 degrees off axis; The test will cover the direction range of 0 deg and ± 20 deg off axis.
– The validation of the design is achieved by checking that the stresses and deformations occurred during the shock loads are within material limits and that the container remains leak tight before, during and after landing on Earth surface (before and after the shock test);
– A proceeding for performing the shock test will be elaborated and information regarding the development of the test bench will be provided.
Functional test with dust contamination over the sealing surfaces: –
SR4, SR5, SR6
– During the sealing of the sample canister the functional test will consider the characteristics of the robotic arm (40 N pushing force, 50 Nm torque);
– The test plan shall provide a phase to prove the capability of the sealing system to properly function after exposure to particle contamination of sealing surfaces;
– The test plan shall contain repeatability phases in terms of sampling, closing and locking;
– Closing and locking operations will be repeated for an acceptable number before and after the tests considering the robot arm characteristics.
– The test plan shall consider using a simulant for the regolith material;
– The test plan shall take into account the particles tendency to adhere on the surfaces;
– The test plan shall include decontamination attempts if the sealing system is provided with a decontamination system.
Functional test taking into account the available robotic arm characteristics SR8:
– The test plan shall include the maximum force (40 N pushing force) and torque (50 Nm) applied by the robotic arm as inputs for developing the test benches.
– If the sealing system requires additional power for closing/locking, the test will outline the necessary force/torque to perform this operation and if it’s within a predefined limit (for example a pushing force of 60 N and a torque of 70 Nm is required) the testing plan will go on in parallel with the design optimisation of the closing/locking mechanism.
Thermal vacuum cycling: –
– Phobos landing site: +27 °C to -143 °C.
– Temperature range non-operational: -40 °C to +40 °C, for a short duration +80°C.
– Test plan shall include well defined temperatures intervals and periods of time based on which thermal cycles can be defined.
FUNCTIONAL ANALYSIS AND REQUIREMENTS ALLOCATION: –
In respect with the previously defined test plan requirements a comprehensive test approach is defined taking into consideration the following aspects:
– objectives (requirements);
– test methods;
– necessary equipment/facilities;
– relevant environmental conditions or test loads;
– test object configurations;
– test object function to be investigated;
– safety requirements when handling the test object;
– details regarding physical or electrical parameters measured during testing;
– criteria defining success or failure of a test.
A test sequence is defined with detailed step-by-step instructions on what the necessary test equipment is, what parameters are to be measured and how the test will be performed.
Designing the ERC
In this chapter, we will study briefly some of the aspects of the design and the capabilities and working functions of the ERC and the vault together. It is essential for us in order for the mission to succeed in our mission, to define the major set of objectives and purposes of the ERC and the vault to resonate parallel to those of the original full mission. Hence, we will describe the objectives and main purposes of them as follows: –
ERC mission: –
• Fully passive re-entry.
• Velocity (w.r.t rotating atmosphere) at EIP < 12,5 km/s.
• Impact of the ERC on the ground.
• Shock load propagation into the internal structure.
• Breaking of the mechanical fuses & releasing of the sample
• Progressive deceleration of the container into the crushable material.
The details regarding the design are coherent and assist in achieving the major goals of the mission and the objectives of the vault. We will describe the major factors that affect the design and how we would go about tackling them.
The proposed landing sites after re-entry are Kazakhstan or Woomera, it will use a full passive re-entry without using a parachute, with a maximum sample container temperature < 40°C and a maximum G-load at landing < 2000g and will posses a radio frequency beacon that will emit radio wave signals for a duration of 4 hours to precisely locate the vault after it has landed. These features are met due to the fact that the crushable and impact absorbent structure enables chute-less landing by protecting the container against landing shock within a prescribed deceleration level