Final Report Summary - AEROFAST (AEROcapture for Future spAce tranSporTation)
Executive Summary:
AEROFAST is a Mars aero-capture feasibility demonstration performed by twelve European companies lead by AST-SAS as prime, and funded under seventh framework programme of the European Commission. This study planned over 2.5 years will end in June 2011.
An aero-capture (A/C) is a flight manoeuvre that takes place at very high speeds within a planet's atmosphere that provides a change in velocity using aerodynamic forces (in contrast to propulsive thrust) for orbit insertion. This aero-breaking technology becomes really attractive with respect to propulsion technology when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions.
Aero-capture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aero-thermal environments, thermal protection systems (TPS), guidance, navigation and control (GN&C), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aero-capture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aero-capture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aero-capture technology up to TRL4 through a complete mission study of a Martian aero-capture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aero-capture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype the thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aero-capture phase recovery.
After two years of studies, the overall mission has been defined and the initial conditions required to perform each phase (cruise, pre-A/C, A/C, post-A/C) established. Mission and systems studies have contributed to generate specifications towards the sub-systems (GNC, communication, power, etc...). The spacecraft architecture has been optimized accordingly leading to a coherent design wrt all disciplines.
GNC algorithms have been implemented within a simulator for NRT (non real time) and RT (real time) tests within a laboratory environment and spacecraft performance has been successfully assessed through an end-to-end mission simulation.
Therefore, AEROFAST study confirms the feasibility of an aerocapture manoeuvre:
* Mission is fulfilled even when considering the worst cases (robust GNC strategy)
* Budgets have been established (mass, power...) coherent each others
No show stopper has been identified, demonstrating the interest of an aerocapture manoeuvre.
In addition, this collaboration of the 2.5 years between all the partners of the project has allowed to build a very good team, in a good spirit, efficient and acting in the interest of the project.
Project Context and Objectives:
1. Introduction
AEROFAST is a Mars aerocapture feasibility demonstration performed by twelve European companies leaded by AST-SAS as prime, and funded under seventh framework program of the European Commission.
This study planned over 2.5 years, started beginning of 2009 and ended in June 2011, for a total budget of about ~3M€.
2. Acronyms
A/C AeroCapture TPS Thermal Protection System
AST ASTRIUM TRL Technology Readiness Level
GNC Guidance, Navigation, Control wrt with regards to
NRT, RT Non Real Time, Real Time
3. AEROFAST: AEROcapture for Future spAce tranSporTation
Aerocapture is a new technology for Solar System exploration that uses a single pass through a planetary atmosphere to decelerate the spacecraft and achieve a targeted orbit (Fig-1).
Such manoeuvre saves a significant amount of mass with regard to a more conventional technique of insertion using propelled braking, and becomes really attractive when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions (sample return missions, in-situ missions and future manned missions which require spacecraft to enter and manoeuvre in a planet's atmosphere).
Figure-1: Aerocapture principle
Aerocapture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), guidance - navigation and control (GNC), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aerocapture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aerocapture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aerocapture technology up to TRL4 through a complete mission study of a Martian aerocapture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aerocapture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype a thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aerocapture phase recovery.
4. Statement of work
4.1 Industrial organisation
The AEROFAST consortium, lead by ASTRIUM-ST-SAS as prime, is composed of twelve European companies from eight countries:
* ASTRIUM-ST SAS (France)
* ASTRIUM-ST Gmbh (Germany)
* DEIMOS Enhgaria (Portugal)
* Amorim Cork Composites S.A. (Portugal)
* SAMTECH (Belgium)
* University of Roma - La Sapienza (Italy)
* STIL-BAS (Bulgarian Academy of Sciences) (Bulgaria)
* Institute of Aviation (Poland)
* Space Research Centre - Polish Academy of Sciences (Poland)
* ONERA (France)
* KYBERTEC (Czech Republic)
* Faculty of Sciences of the University of Lisbon (Portugal)
4.2 Work breakdown structure
The work breakdown structure is shared out of five work packages in order to fulfil the objectives of the project, and a sixth work package with deals with management, dissemination and exploitation. Details are reported Tab-1.
Table-1: WP breakdown
The work package leaders are:
WP1: AST-ST SAS WP4: DEIMOS Enhgaria
WP2: AST-ST GmbH WP5: University of Roma
WP3: Amorim Cork Composite WP6: Kybertech and AST-ST SAS
4.3 List of deliverables
The deliverables have been released according to the deliverable item list reported next Tab-2. These documents are available on the EC FP7 server.
Table-2: Deliverables items list
4.4 Planning
The entire milestone has been passed according to the schedule.
The final presentation has been done in June, 28th 2011. The last deliverables have been sent end of August 2011, as agreed by the EC. From the technical point of view the AEROFAST contract was ended at T0 + 30 months. From the contractual point of view the contract is finished at T0 + 33 months.
Figure-2: planning of AEROFAST study
Project Results:
1. Introduction
AEROFAST is a Mars aerocapture feasibility demonstration performed by twelve European companies leaded by AST-SAS as prime, and funded under seventh framework program of the European Commission.
This study planned over 2.5 years, started beginning of 2009 and ended in September 2011, for a total budget of about ~3ME.
2. Acronyms
A/C AeroCapture IMU Inertia Measurement Unit
AGL Above Ground Level L/D Lift-over-Drag
AST ASTRIUM LGA Low Gain Antenna
BC Back Composite LOS Line of Sight
CAC Cruise & Aerocapture Composite MO Mars Orbiter
DOF Degree of Freedom NRT, RT Non Real Time, Real Time
DSN Deep Space Network OBDH OnBoard Data Handling
EIP Entry Interface Point P/L PayLoads
FC Front Composite TCM Trajectory Change Manoeuvre
FPA Flight Path Angle TPS Thermal Protection System
GNC Guidance, Navigation, Control TRL Technology Readiness Level
HGA High Gain Antenna wrt with regards to
IMS Imaging Multi Spectral
3. AEROFAST: AEROcapture for Future spAce tranSporTation
Aerocapture is a new technology for Solar System exploration that uses a single pass through a planetary atmosphere to decelerate the spacecraft and achieve a targeted orbit (Fig-1).
Such manoeuvre saves a significant amount of mass with regard to a more conventional technique of insertion using propelled braking, and becomes really attractive when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions (sample return missions, in-situ missions and future manned missions which require spacecraft to enter and manoeuvre in a planet's atmosphere).
Figure-1: Aerocapture principle
Aerocapture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), guidance - navigation and control (GNC), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aerocapture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aerocapture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aerocapture technology up to TRL4 through a complete mission study of a Martian aerocapture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aerocapture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype a thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aerocapture phase recovery.
4. Main technical achievements
4.1 Mission overview
The frame of the AEROFAST mission has been defined as a low cost mission. The launch and transfer orbit injection is supposed to be performed by a Soyuz Fregat departing from Kourou, and able to perform the coast phase needed to get rid of declination limitations.
The complete mission sequence contains the phases of:
* Cruise & Pre-aerocapture phase: Challenge is to master the attitude/position of the spacecraft (S/C) before manoeuvre:
- Cruise : Several trajectory correction manoeuvres (TCM) are foreseen to allow Martian rendez-vous with the use of star trackers for navigation
- Approach: Before arriving to Mars, the Deep Space Network (DSN) navigation will be hybridized with an optical + radiometric navigation increasing the accuracy of targeting the aerocapture corridor.
* Aerocapture phase: Used to brake the incoming velocity thanks to the drag exerted by the atmosphere and will insert the spacecraft on the appropriate elliptical orbit.
* Post aerocapture phase: to target a quasi circular sun-synchronous orbit at a low altitude.
- Science Orbit Acquisition: Before reaching its operational orbit, the spacecraft will perform some transition manoeuvres that, with the help of the drift due to gravitational perturbations, will drive it to its final orbit.
- Science Orbit Operation: Once inserted in its final orbit, the vehicle will begin its observation mission (swath width of 25km with a resolution of ~1m) and will only perform small orbit adjustment manoeuvres.
4.2 Mission analysis
4.2.1 Mission constraints
Soyuz from Kourou offers a range of payload compatible with AEROFAST type of mission, and the flexibility to reach different declination with the re-ignition capability of Fregat upper stage.
From the mission analysis, a set of constraints reported Tab-3 has been identified which drives the cruise and pre aerocapture phase.
Constraints Justification
RQ-1 Soyuz performance must be greater than 1450 kg (1250kg spacecraft + 200kg ACU) for the launch opportunity (infinite velocity and declination at departure
RQ-2 Infinite velocity at arrival must be lower than 4 km/s To avoid energy dissipation issue during aerocapture (aerocapture design constraints)
RQ-3 The angle between the infinite velocity and the Earth direction must be lower than 90° at Mars Arrival To have aerocapture phase placed on the side facing the Earth, with the communication link available (even if Mars is hidden by the Sun at space craft arrival), and with operational transmitters during aerocapture phase (pre-aerocapture communication constraints)
RQ-4 The Sun-Earth-Probe (SEP) angle must be greater than 7° from 20 days before Mars arrival
RQ-5 The angle between the Earth direction and the vehicle velocity has to be greater than 40° during the whole aerocapture
RQ-6 The angle between the infinite velocity at arrival and the Sun direction must be greater than 45° To be able to use optical measurements (stars trackers, camera) on Mars moon in the last 20 days before re-entry (pre-aerocapture navigation constraints)
RQ-7 Pre-aerocapture navigation performance and propagation of S/C release down to EIP must be such that FPA errors at entry point are below ±1.24 deg. To allow re-entry within the aerocapture corridor (aerocapture guidance constraints)
Table-3: Constraints from mission analysis
4.2.2 Cruise & Pre-aerocapture phase
The analysis of the Earth to Mars opportunities has been performed for the whole decade 2020-2030. Considering the constraints reported Tab-3, a set of 21 opportunities has been identified (Fig-3).
The best launch opportunity from Earth has been found in Mai, 11th of 2026, with a Mars arrival expected in June, 11th of 2028 as the best compromise between minimum departure infinite velocity and transfer duration from Earth to Mars (opportunity n°9 among 21).
Figure-3: Constraint compliance map (white = no compliance)
Figure-4: AEROFAST transfer from Earth to Mars
The features of the selected trajectory are:
* A transfer of type T4 which suppose to have more than one revolution about the sun, with 6 TCM to meet the target conditions at Mars arrival (Fig-4). TCM have been optimised wrt the navigation process and the overall delta-V required to be minimized. Last TCM is foreseen not later than EIP -4h
* An injection into polar orbit at Mars by any of the two possible arrival branches (north and southbound)
* An Entry Interface Point (EIP) defined at 120km AGL, with a Flight Path Angle (FPA) of -10.74 deg.
Zone Transfer type Min. depart. velocity (km/s) Declination
(deg) Departure date Arrival date Transfer time (months)
1 T1/T2 4.3057 -10.786 31/07/2020 14/07/2021 11.43
2 T1/T2 3.7219 14.880 15/09/2022 30/09/2023 12.48
3 T1/T2 3.3342 19.735 04/10/2024 13/09/2025 11.30
4 T1/T2 3.1070 32.894 05/11/2026 13/10/2027 11.24
5 T1/T2 3.0002 28.250 10/12/2028 05/11/2029 10.84
6 T4 2.9756 29.130 05/12/2021 02/03/2024 26.87
7 T4 2.9811 19.921 04/01/2024 04/03/2026 25.95
8 T4 3.2276 10.642 29/01/2026 20/02/2028 24.71
9 T4 2.7816 -14.641 11/05/2026 11/06/2028 25.03
10 T4 3.5997 19.114 21/09/2028 27/04/2031 31.15
11 T5 3.1335 -49.399 19/04/2021 29/05/2024 37.32
15 T6 3.4743 6.046 16/02/2021 10/06/2024 39.75
16 T6 3.3289 21.477 06/10/2021 19/09/2025 47.44
17 T6 4.6517 -5.999 27/01/2023 15/04/2026 38.57
18 T6 3.2223 29.179 20/10/2023 17/09/2027 46.92
20 T6 3.1869 32.772 08/11/2025 18/09/2029 46.32
Table-4: AEROFAST selected trajectory
4.2.3 Aerocapture phase
The aerocapture feasibility is mainly driven by the conditions at the EIP (Tab-5). The aerocapture corridor is bounded by extreme over-shoot (hyperbolic exit) and under-shoot (crash) trajectories. The small width of the corridor leads to maximum FPA errors bellow ±1.24 deg at EIP.
Entry Interface Point (EIP) at 120km AGL
Relative velocity: 6377.6 m/s
Flight Path Angle: (local geodetic frame) -10.74°
Azimuth of relative velocity: -19.59°
Latitude: -77.8°
Longitude: (geodetic frame) 10,6°
Table-5: EIP features
Figure-5: Entry corridor for Aerocapture at Mars arrival
As depicted Fig-6, the aerocapture phase allows targeting the orbit inclination while exiting the atmospheric phase, but aerocapture doesn't allow meeting all in plane parameters, in particular pericenter altitude of the orbit for science operation (Tab-6).
At the end of the aerocapture at 120km altitude, spacecraft exit parameters (relative velocity ~3500m/s, FPA ~3°) lead to the nominal apocenter altitude (345km AGL) but a too low pericenter altitude (-44km ->crash).
Thus, a fast correction manoeuvre must be performed just after aerocapture to reach the nominal pericenter altitude required (290km AGL).
4.2.4 Post Aerocapture phase
The orbit for sciences operations must comply with the requirements of the scientific optical payloads. AEROFAST orbit reported Tab-6 has been chosen in order to have an almost global coverage with repeating ground track and constant lighting conditions:
* Repeating orbit condition, which allow to have image at the same location with similar light incidence condition
* Maximum surface coverage, which allow to observe the largest part of Mars
* Sun synchronous condition, which allow to compensate the Mars motion around the Sun to keep same lighting conditions
* Frozen orbit condition, which correspond to an orbit with parameters having small rates of variation in order to reduce the cost of the orbit control
Semimajor axis 3713,87 km Pericenter altitude 290.55 km
Eccentricity 7,355e-03 Apocenter altitude 345.18km
Inclination 92,7° Orbital Period (OP) 6871s = 1.91 h
Pericenter argument 270° Mars global coverage 1098 OP
Table-6: orbit features for sciences operations
4.3 Guidance, Navigation and Control architecture
The GNC architecture is composed of sensors, actuators and a guidance & control philosophy to meet the navigation requirements of each mission phase. The reference used for GNC studies is the baseline spacecraft (shape, mass) reported §6.
4.3.1 Cruise & Pre Aerocapture phase
During the cruise phase, the spacecraft is in sun pointing mode. The only changes in spacecraft orientation are for the 6 TCM and the final target orientation.
Spacecraft navigation relies on a class-3 Inertia Measurement Unit (IMU) and star trackers during cruise phase coupled with a deep space network transponder. Sun sensors may also be activated in a safe mode.
Control is done by reaction inertia wheels (for attitude control) and thrusters (for TCM and inertia wheels desaturation). For Mars final approach (EIP-5days), the optical camera is pointed towards Mars (LOS measurements) and radiometric measurements are activated to meet the requirements at EIP.
Monte Carlo simulations, including external perturbations and sensors performances, have demonstrated the robustness of the GNC architecture with accuracy < 3km (3σ) at EIP for a total delta-V < 30m/s for pre-A/C phase.
4.3.2 Aerocapture phase
Aerocapture phase must be safe (no crash or hyperbolic exit) while being complaint with the mission constraints (pre-A/C) and requirements (post-A/C). The philosophy consists in decoupling in-plane & out-of-plane motions for respectively apoapsis and inclination control, while remaining as much as possible in the aerocapture corridor (semi permeable domain).
Figure-6: In-plane & out-of-plane required motions
Figure-7: Nominal guided trajectory (not sizing)
Spacecraft navigation during aerocapture relies on the IMU (maximum deceleration about ~4g in worst case), whereas Control is performed by a set of thrusters which maximum torques have been assessed at 200N.m around the yaw axis, 100N.m around pitch and roll axis, with a minimum impulse bit (MIB) < 25ms.
The guidance performance has been assessed by Monte Carlo simulation (1000 runs, 3DOF). Results show no crash, no hyperbolic exit, an inclination well controlled and a 100% fulfilment rate of the total delta-V < 170m/s (Fig-8) for aerocapture phase and science orbit acquisition.
Figure-8: Aerocapture robustness (MtC simulations)
The performance of the control has been assessed (Fig-9) for the nominal trajectory leading to oscillations amplitude below ±5° for both angle of attack and sideslip angle (< ±1° for Pdyn >100Pa), and bank angle manoeuvres performed with accuracy better than 5°.
Figure-9: Control performance during nominal guided trajectory
4.3.3 Post Aerocapture phase
Just after aerocapture, TPS heat shield has to be jettisoned to allow fast correction manoeuvres to target the in-plane parameters of the science operation orbit:
* During the science orbit acquisition, navigation relies on IMU. The two burns which are required at apocenter & pericenter (mainly to shift pericenter altitude from -44km to 290km) are performed by the attitude control thrusters of the Mars Orbiter.
* Once Mars Orbiter placed on the science operation orbit, navigation relies on the star trackers and the IMU. The attitude control is performed by reaction inertia wheels, while the slow decrease of the semi-major axis (Fig-10) is corrected by bi-impulsive corrections done by the thrusters every 61 orbits to limit ground track drift at 30m maximum at the equator wrt the nominal ground track.
The science orbit control requires a delta-V~3.6m/s for 1 period of repetition (1098 nodal periods) corresponding to a full coverage of the Mars surface.
Figure-10: Evolution of semi major axis (if no correction) compared with the stability of the argument of perigee over an entire period of repetition of 1098 nodal orbital periods (frozen orbit conditions)
4.4 Spacecraft design
Missions with aerocapture phase generally involve spacecraft architecture based on two main modules, an Orbiter module encapsulated within a protective module jettisoned after Aerocapture.
Figure-11: Product tree
AEROFAST spacecraft product tree depicted Fig-11 is based on three different modules, a Mars Orbiter, a Cruise and Aerocapture Composite composed of a Front Composite and a Back Composite (option):
* During pre-aerocapture and aerocapture phase, the Mars Orbiter module is implemented within the Cruise and Aerocapture Composite (Fig-12)
* After aerocapture, prior to post-aerocapture manoeuvres, Front & Back Composites are jettisoned (Fig-12)
* Post-aerocapture manoeuvres for sciences orbit acquisition & operations are performed by the Mars Orbiter (Fig-12)
Figure-12: Aerofast different configurations
4.4.1 Aero-shape trade-off
Aerocapture requires an aeroshape that provides a lift-over-drag (L/D) ratio with sufficient provision (L/D>0.3) and sufficient static stability performances. In addition, the aeroshape must protect the payloads (P/L) from the severe aerothermal environment during the aerocapture, and shall be safely jettison able at the beginning of the post-A/C phase.
In the frame of AEROFAST, a trade-off has been performed considering two configurations:
* A blunt body capsule derived from Apollo, ARD concepts
* A lifting body dealing with bi-conic shape
By considering L/D requirements, payload protection wrt aerothermal heating and layout constraints, the bi-conic shape has been selected.
4.4.2 Aerothermodynamics analysis of the biconic shape
Aerodynamic performances of the bi-conic shape have been assessed for different angle of attack by Euler calculations, results exhibiting L/D ratio > 0.4 at angle of attack < 40deg (Fig-13).
Figure-13: Max dynamic pressure point - Mach 27
Figure-14: Satellite flow interaction t=91s, 40° angle
Spacecraft-flow interactions (Fig-14) have been analysed, demonstrating (Fig-15) the need of a Back Composite protection for the Orbiter (locally heat fluxes >20kW/m2 mainly due to flow radiation)
Figure-15: Radiation heat flux on Orbiter at t=91s
Figure-16: Max heat flux history (T°w=300K)
Heat fluxes histories (Fig-15) have been assessed along the two worst case aerocapture trajectories identified by Monte Carlo analysis:
* overshoot trajectory corresponding to the maximum heat flux trajectory, with a maximum flux ~1064KW/m2 found at stagnation point (turbulent regime applied over the whole surface wrt Reynolds criteria)
* undershoot trajectory corresponding to the maximum energy trajectory
The baseline TPS foreseen to protect the spacecraft is the Norcoat liège material (used on ARD, Beagles), glued on Front & Back Composite structures. The TPS thickness is driven by maximum temperature at the interface with the structure (T°max ~450K driven by the silicone based glue).
3D computations results considering Norcoat pyrolysis/charring and ablation lead to:
* thickness from 19mm to 6mm over the Front Composite surface, with a maximum ablation of 10mm
* thickness of 3.3mm uniform over the Back Composite surface
4.4.3 Spacecraft architecture
Spacecraft architecture results from trades-off performed at system level, with permanent drivers for saving mass and increasing robustness by selecting principles as simple as possible.
A trade-off has been conducted about the Cruise Stage. Several options were assessed: 2 configurations with external Cruise Stages jettisoned before aerocapture and 1 configuration with an integrated Cruise Stage. At last, the integrated Cruise Stage has been selected has the best compromise between mass saving and robustness (no additional separation mechanism).
The baseline architecture of the spacecraft (CAC and MO) is depicted Fig-17.
Figure-17: Baseline architecture of AEROFAST CAC and MO
The design and the layout of the spacecraft are driven by the mission sequence:
* Spacecraft must comply with the launch vehicle fairing and interplanetary orbit injection mass capability (Soyuz from Kourou)
* Power generation & storage, communication link to Earth must be available during cruise phase and science orbit acquisition & operation phases
* Autonomous guidance, navigation and control must be available along the entire mission from cruise, approach, aerocapture, post aerocapture phases.
A preliminary thermal control analysis has been performed which concluded that no active thermal control system is required for on board equipments. However, a cold case has been identified on hydrazine tanks which may require active heater during cruise.
The functional architecture of the spacecraft depicted Fig-19 has been established with the desire to make re-use of components as much as possible, and to fulfil redundancy requirements.
4.4.4 Spacecraft layout
Most parts of the power, On-Board Data handling (OBDH) and communication systems are integrated within the MO. The equipments which are solely needed during cruise and aerocapture phases are installed within the CAC. The equipments required for the entire mission which need to have outside access have been duplicated and implemented both within CAC and MO (solars array, communication antenna and attitude thrusters).
The specifications of the onboard equipments (measurement accuracy, torques, power budget etc...) have been established wrt the system studies, and by considering the worst cases identified by the Monte Carlo analyses.
Attitude & orbit control subsystem
Navigation:
* Star trackers and sun sensors both on CAC and MO
* IMU (NG LS-200S) for aerocapture and orbit acquisition within MO
* Navigation camera for approach phase on CAC
Control:
* Reaction wheels for cruise and orbital phase within MO
* Thrusters:
- 6 aerojets MR120 (90N) on CAC for aerocapture + 4 CHT-5 (5N) for reaction wheels desaturation during cruise
- 12 CHT-20 (20N) on MO for science orbit acquisition, attitude control, and inertia wheels desaturation
* 1 tank of Hydrazine (N2H4) is required within CAC to cover cruise, pre-A/C and A/C needs (113kg propellant), whereas 2 tanks are required within MO to cover post-A/C needs (~28kg)
Figure-18: Equipments considered
Figure-19: AEROFAST functional architecture
Power subsystem
The power architecture is based on a hybrid approach, a Direct-Energy-Transfer (DET) for cruise phase which is quasi permanently sun illuminated and a Peak-Power-Tracking (PPT) for orbital phase which is dominated by alternative eclipses/daylights in short period:
* During cruise & pre-A/C phases, power needs are in average ~115W which are provided by solar cells of ~1,23m2 fixed on the CAC surface.
* For orbital phase, power needs vary from 143 watts (eclipse) to 562W (needs for observation and transmission during daylight) provided by 2 solar arrays of 5,6m2 each on MO.
Batteries (8x VES 140 cells) are implemented within the MO with a total capacity of 39A/h at 28,8V.
Figure-20: Batteries and capacities
Communication subsystem
The communication subsystem provides a data link for housekeeping, telemetry/telecommand and transmission of scientific data.
* 4 Low Gain Antennas (LGA) are used for telemetry/telecommand during cruise and for orbit acquisition phases. LGA are implemented both within CAC and MO, and in a redundant way (mission critical).
* 1 High Gain Antenna (HGA) is used for transmitting scientific data to Earth. Antenna diameter is foreseen to be 2.1m diameter, with a transponder operating in Ka-band at minimum data rate of 594kbit/s. HGA is implemented within MO.
Scientific payload
Science payload is assumed to be 125kg maximum, with a maximum power consummation of 100W.
For the moment, an Imaging Multi Spectral (IMS) sensor is foreseen (87,5kg, 55Watts) dedicated to take pictures for a global observation of Mars surface (panchromatic scanner: 500-800 nm; multispectral scanner: 3 channels: Green, Red, NIR; Spatial resolution, IFOV: 1m (Pan); 4m (Multi) in NADIR position; swath ~23km)
Instrumentation is also foreseen for environment characterisation and TPS analysis during aerocapture phase (see §5.6).
4.4.5 Mass breakdown and geometry
Mass breakdown, geometry and centring are depicted Fig-21 and Tab-7.
Table-7: AEROFAST mass breakdown
Figure-21: CAC shape and centring box
4.5 Validation of the GNC performances
4.5.1 Methodology
Based on the mission, architecture and general design of the spacecraft, the GNC algorithms performance for Pre-A/C and A/C phases has been assessed through 6DOF simulators with Non Real Time (NRT) and Real Time (RT) breadboards.
Both pre aerocapture and aerocapture performances have been assessed separately. The logic is depicted next Fig-22.
Figure-22: development flow
As explained previous sections, GNC architecture relies on both optical (1 camera) and radiometric measurement activated 5 days before EIP, last correction manoeuvres being feasible until 4h before EIP.
4.5.1 Pre aerocapture GNC assessment
4.5.1.1 Physical mock up (IP-lab)
Optical navigation is a subsystem to GNC that can provide information on spacecraft (S/C) location and attitude with respect to a celestial body. In AEROFAST context, this applies to the Pre-Aerocapture mission phase. Main optical observables are the Line-of-Sight (LOS) to a target and image dimensions (diameters for spherical bodies).
To verify the performance of such observables it is necessary to recreate (within a valuable scope) the imaging conditions. Therefore, the IP-lab mock-up depicted Fig-23 has for objective to simulate the mission profile with real images in order to test imaging and GNC combined performances.
Images are generated by a live camera looking at a ball lit by a source with adequate irradiance. The produced image does simulate the solar illumination and the range of Mars planet.
Figure-23: IP lab (Mars/Sun unit)
The calibration and test campaign has allowed validating the IP algorithms in relevant cases with real camera performances.
4.5.1.2 Pre aerocapture Non Real-Time GNC performance (FES)
NRT Monte Carlo simulations (200 runs) for Pre-A/C phase include sensibilities to initial position and velocity errors at EIP-5days, as well as to mass, propulsion and relative biases between camera and attitude sensors.
Results reported Tab-8 show that conditions at EIP are met even by considering worst case (initial position and velocity errors, increased misalignments) which validates pre-A/C GNC architecture. The simulations indicate that relative bias is the most important contributor to the GNC performance. A relative bias of 72µrad (3σ) leads to comply with the requirement at EIP of 3km (3s), a contrary to a bias of 290µrad (3σ).
Figure-24: Pre Aerocapture NRT simulation (FES) and test cases selected
Figure-25: Spacecraft position dispersion wrt target position at EIP (XY and XZ view)
(MC2 test case)
Figure-26: Statistic on pre-A/C position and velocity errors, FPA dispersions wrt target FPA (MC3 test case)
Table-8: Pre A/C NRT tests results
4.5.1.3 Pre aerocapture Real-Time GNC performance (HIL)
The Real Time simulation is performed to prove the GNC capability to run on a LEON3 processor in a nominal way. The simulation starts 5 days before EIP. Two cases have been selected from the NRT Monte Carlo simulations: Nominal and Robustness case.
Figure-27: Pre A/C RT tests results (Robustness case)
Requirements are fulfilled and results are coherent of the NRT simulations for both cases, demonstrating the feasibility & robustness of the autonomous pre A/C GNC concept.
4.5.2 Aerocapture GNC assessment
GNC design for aerocapture is based on:
* For Guidance: a numerical apoapsis predictor-corrector derived from MSR-O and ATPE studies
* For Navigation: A5-like inertial navigation with autonomous alignment (class-3 IMU)
* For Control: a flight proven ARD-like algorithm
4.5.2.1 Aerocapture Non Real-Time GNC performance (FES)
NRT Monte Carlo simulations (6DOF, 500 runs each) have been performed for A/C phase in order to assess GNC performance and robustness by considering a set of dispersion (EIP, atmosphere, aerodynamic coefficients, MCI, IMU).
Results reported Tab-9 shows 3 violations of the requirements.
* Dynamic pressure (~6.4Kpa instead of 6Kpa max)
* AoA at low pressure (5.4° instead of 5° max)
* AoA at high pressure (1.8° instead of 1° max)
But these violations which occur for some dispersed cases are very slight and not considered as critical.
Table-9: A/C NRT tests results (Monte Carlo 6DOF, 500runs)
4.5.2.2 Aerocapture Real-Time GNC performance (PIL)
The Real Time simulation is performed to prove the GNC capability to run on a LEON3 processor in a nominal way as for Pre-A/C phase, and to check the GNC needs wrt the available resources (CPU, delays, memory) as dynamic phenomenon are highest during the A/C compared to the pre-A/C phase.
The RT results reported Tab-10 are in line with the NRT A/C simulation results.
Table-10: A/C RT tests results (nominal)
GNC Real Time results are good even if simulations have revealed impact of the GNC computational delays, mainly guidance computation time (up to 0.8 s). Maximum CPU load vary from 65 % (nominal) to 90 % (worst-case)
4.6 Instrumentation
A preliminary instrumentation plan has been established in order to cover the need of data for three different objectives: environment analysis, TPS analysis and scientific payloads.
For each proposed sensors, compatibility with the environment and performance wrt the requirement has been assessed. In addition, a functional architecture has been proposed based on a sensor network built on two levels: measuring cells, containing different types of sensors and cluster controller, collecting measurements and communicating with the on board computer.
4.6.1 Environment analysis sensors
The sensors are dedicated to provide data about the environment during the re-entry phase. The two categories of sensors foreseen are integrated pressure and heat flux sensor (Raflex type) and temperature sensors implemented within thermo plugs.
The Raflex type sensor is used to measure both pressure and heat flux. It is composed of a pressure tap and a calorimeter made of Molybdenum as depicted Fig-28a. Thermo plugs are made of 3 thermocouples each placed at different depths (Fig-28b).
Figure-28 : RAFLEX sensor (a), thermo plug (b)
5 Raflex sensors and 16 thermo plugs are foreseen on the front composite.
4.6.2 TPS analysis sensors
The main interest consists in collecting information on the plasma state and the chemical composition of the TPS surface during the aerocapture phase. Spectrometric measurements are foreseen by using pyro-sensors & several filters as depicted Fig-29.
Figure-29: Pyro sensor
5 pyro sensors are foreseen according to 5 different generatrices.
4.6.3 Scientific payloads
The goal of the mission after the aerocapture phase is the global observation of the Martian surface by pictures taken by an optical payload and transmitted to Earth.
The selected payload is an Imaging Multispectral Sensor (IMS) whose design and performances are depicted Fig-30.
Figure-30 : Scientific payload (IMS)
4.7 TPS improvments
Norcoat liege TPS (cork based) has been chosen as the baseline for Aerofast mission wrt the demonstrated performances of this TPS for atmospheric entry probes (ARD, Beagle-2).
Nevertheless, research and development of advanced cork based TPS have been initiated in order to improve mass, performance, resistance to erosion, resistance of the char material and surface aspect after degradation. The objective is to propose reinforced material for front shield and superlight material for back shield. Several tracks have been followed including:
* different cork granules
* different types of resins (phenolic, furan)
* different types of fillers (carbon fibres, basalt fibres, glass spheres)
* optimisation of mixing process
Preliminary basic characterisations leaded to select the most promising formulations for complementary plasma tests in order to assess thermal performance and ablation, for comparison with Norcoat liege & P50 reference materials.
Two mission profiles have been selected corresponding to the maximum energy and heat fluxes trajectories. Samples have been manufactured for plasma tests (Comete facility).
Figure-31: selected material for plasma tests campaign
Globally, behaviour of all material is quite good, with better surface aspect for both mission profiles. Masse losses are of same order of magnitude, but for few samples, thermal insulations seems not so good (those with carbon fibres). Best results have been found with the TPS3J formulation (phenolic + carbon fibres)
Figure-32: Plasma tests on Cork based material
Therefore, tests demonstrated that reinforcement with fibres improves the resistance to ablation, confirmed that phenolic grade resin is still the most appropriate compared with new resins tested (green resins).
The manufacturing of a very complex prototype (small radius, thickness not constant as depicted Fig-33) demonstrated that cork based TPS offer wide possibilities for moulding or machining panels, but expertises have shown important density heterogeneities. Therefore, complex shapes are achievable with cork based TPS, but manufacturing process has to be optimised accordingly.
Figure-33: TPS prototype
A non ablative option has been also assessed. A preliminary design based on C/SiC TPS for FC and FEI TPS for BC leads to a total TPS mass of about ~150kg, which is similar to the TPS mass budget with Norcoat liege. But, heat-shield robustness and maturity based on C/SiC TPS is lower compare with cork based TPS. C/SiC could be interesting only if FC underlying structure could be removed (mass saving ~75kg).
5. Assessment of the objectives
AEROFAST project has been established wrt five top objectives
* OBJ-1: Define a project of aerocapture demonstration -> completed
- Mission analysis has been performed -> set of requirement for system studies (ATD, Aero, GNC) -> set of requirements for sub systems (spacecraft design, propulsion, power etc...)
- A coherent project of an aerocapture has been defined (mission, system, sub-system)
* OBJ-2: Increase TRL up to 4/5 of the planetary relative navigation and the aerocapture algorithm -> completed
- TRL5: breadboard validation within a representative environment.
- Algorithms for pre-aerocapture and aerocapture phases have been developed according to the spacecraft design, and performance has been assessed through breadboard simulations (NRT, RT)
* OBJ-3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms -> partially completed
- Objective closely linked to the objective-2. Simulations in RT of the GNC algorithms and the hardware (IP-lab) have been successfully validated but separately.
* OBJ-4: Demonstrate/prototype the TPS for the mission -> completed
- Usual cork based TPS is suitable for an aerocapture mission
- Non ablative TPS comparable in term of mass to the ablative material
- Improvement of cork based material has been performed (roughness improvement of the charred surface) with the addition of fibres
* OBJ-5: Define O/B instrumentation for aerocapture phase recovery -> completed
- On-board instrumentation has been defined for: navigation, TPS/environment characterisation and scientific observations
6. Conclusions
AEROFAST is not dedicated to perform a phase-A, but to prepare for such a mission demonstration by increasing technology maturity to TRL4 (mainly GNC & thermal aspects).
After two years of studies, the overall mission has been defined and the conditions (EIP...) required to perform each phase (cruise, pre-A/C, A/C, post-A/C) established. Mission and systems studies have contributed to generate specifications towards the sub-systems (GNC, communication, power, etc...). The spacecraft architecture has been optimized accordingly leading to a coherent design wrt all disciplines.
GNC algorithms have been implemented within a simulator for NRT and RT tests within a laboratory environment and spacecraft performance has been successfully assessed through an end-to-end mission simulation.
Therefore, AEROFAST study confirms the feasibility of an aerocapture manoeuvre:
* Mission is fulfilled even when considering the worst cases (robust GNC strategy)
* Budgets have been established (mass, power...)
No show stopper has been identified, demonstrating the interest of an aerocapture manœuvre.
In addition, this collaboration of the 2.5 years between all the partners of the project has allowed to build a very good team, in a good spirit, efficient and acting in the interest of the project.
Potential Impact:
From the past experience on aerocapture, this transportation technology is clearly identified as core technology for the upcoming future missions and will save significant mass at launch. The past activities highlighted the issues and the complexity of such a transportation technology. TRL is in the order of 2 to 3 in Europe. In order to use the aerocapture technology for operational missions while mitigating future development risks, the TRL of such a technology must reach the level 6. TRL6 shall be reached through a flight demonstration.
Such a demonstration can only be envisaged at European level and shall be led by ESA, the European Space Agency. Assuming that the overall development of such a mission is in the range 400 to 600 M€, 5% of the budget is necessary for the feasibility assessment (called phase A). The approach taken by EC is to help Agencies to prepare for the future and to fund space technologies supported by Agencies with clear application on future programmes and those Agencies cannot fund (too long term, too risky ...).
So clearly AEROFAST is in line with this approach: AEROFAST end goal is not to perform a phase A of the mission demonstration but is to prepare for such a mission demonstration and increase the TRL level: AEROFAST intends in the frame of the FP7 to reach a TRL 3 to 4 for this technology. Moreover the AEROFAST partners have been chosen in accordance with these goals. The AEROFAST team is not tailored to run a phase A of a mission demonstration. Anyway from the cost point of view the budget dedicated to AEROFAST does not allow building and performing an overall feasibility assessment of the demonstration. It might be précised that the consortium worked very well together and is spread over eight countries. At least these different partners from these different countries would be really enthousiastic to go further on the project and develop it together until flight.
Following tasks were envisaged within AEROFAST to increase the TRL level up to 3 / 4:
First of all a mission were drafted and mission and design requirements were assessed. This provided the framework of AEROFAST and the conditions for the other tasks.
Secondly, AEROFAST worked in parallel on several topics necessary for TRL increase:
* Increase our knowledge on aerothermodynamics and flow behaviour,
* Improve and better assess the effect of ablation with respect to the centre of gravity,
* Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5,
* Demonstrate/prototype the thermal protection system for such a mission.
In addition it might also be precised that the different areas of research of AEROFAST could also be more widely used and results obtained used for other applications. Table n°1 highlights the growth potential of AEROFAST results to other applications and identifies the additional tasks to be implemented. AEROFAST is built and studied for a demonstration in a Martian environment, nevertheless interest is to identify whether some of the disciplines could be adapted to other applications:
* Disciplines linked to the necessity of having an atmosphere, and possibly different from Mars,
* Generic disciplines.
Growth potential to
Disciplines & products: Mars application Another atmosphere generic tasks
Mission
vehicle design:
heat shield
cruise module
orbital module X
X
Valid for the cruise module and for the orbital module - to be adapted to the mission needs
aerocapture corridor X X Methods to compute corridor can be re-used
Corridor in Earth different from the one on Mars
Aerodynamics X CO2 environment with winds & storms on Mars
Very different physical phenomena
Shape of the front shield might be different from one atmosphere to another
Aerothermal Environments X
Innovative thermal Protection System X X Material based on cork does not fit for an aerocapture on earth (due to very high thermal environment)
Results obtained could be assessed in other environmental conditions in order to test their behaviour, for instance an re-entry from the ISS orbit (max 400 km)
Necessity to perform computations and associated ground tests in the facilities
Aerocapture GN&C systems
associated simulator
test bench X X for G
X Guidance algorithms independent from the atmosphere
navigation and control choices pending upon Mars
test bench generic, only sensors to be adapted to the mission to be tested
Disciplines & products: Mars application Another atmosphere generic tasks
On-board navigation systems X X Optical navigation to be adapted to Earth because of the planets sizes
System could be re-used for another planet, no interest around earth (GPS)
Innovative on-board instrumentation X Instrumentation for the aerocapture phase is dedicated to Mars
Table n°1 Utilisation of AEROFAST results to other applications
More globally speaking AEROFAST participates:
* to sustain a competitive industry (including manufacturers, services providers) in defining an aerocapture demonstration on Mars while making a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5, as well as prototyping an innovative thermal protection system for such a mission and innovative on-board instrumentation for aerocapture knowledge improvement,
* To provide appropriate services and infrastructures.
Moreover the project AEROFAST fully agrees
* as well with a progress towards the sustainable provision of technologies needed by the European Space to become non dependent,
* As with the consolidation of long term sustainability,
* as with the improvement of the economical aspects of a domain known to be demanding in terms of reliability, experimenting novel techniques (relative navigation) and methodologies (thermal protection system).
Regarding Space independence, this aerocapture technology will be a potential solution for several future applications:
* Sample return missions,
* Manned missions returning from the Moon,
* Robotic missions towards planets such as Mars, Venus or Titan.
Those potential applications are inline with the European space policy in terms of space transportation and ESA approach regarding future space exploration: through Moon and Mars Missions.
All countries (USA, Europe and Russia) agree on the importance and interest of aerocapture for future high mass cargo missions and manned missions, to allow for human expansion into the solar system. All past studies and projects show that such transportation vehicles must rely on aerocapture to be mass effective: using atmospheric drag to slow transportation vehicles is regarded as one of the largest contributors to making both lunar and Martian missions affordable.
The MSR mission and future manned missions are scheduled not earlier than 2022, so roughly speaking in more than 15 years. On Moon side some early robotic missions are foreseen in order to improve the necessary technologies for MSR, focused on rendezvous and capture and soft-landing - aerocapture here is not addressed. NASA would also envisage manned missions on the Moon around 2020 where aerocapture would become of interest for returning the crew to Earth. Globally speaking Aerocapture transportation technology should be mastered within the next decade.
Thanks to AEROFAST a good step is likely to be achieved and allows for further improvements and knowledge on this technology.
AEROFAST is a Mars aero-capture feasibility demonstration performed by twelve European companies lead by AST-SAS as prime, and funded under seventh framework programme of the European Commission. This study planned over 2.5 years will end in June 2011.
An aero-capture (A/C) is a flight manoeuvre that takes place at very high speeds within a planet's atmosphere that provides a change in velocity using aerodynamic forces (in contrast to propulsive thrust) for orbit insertion. This aero-breaking technology becomes really attractive with respect to propulsion technology when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions.
Aero-capture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aero-thermal environments, thermal protection systems (TPS), guidance, navigation and control (GN&C), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aero-capture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aero-capture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aero-capture technology up to TRL4 through a complete mission study of a Martian aero-capture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aero-capture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype the thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aero-capture phase recovery.
After two years of studies, the overall mission has been defined and the initial conditions required to perform each phase (cruise, pre-A/C, A/C, post-A/C) established. Mission and systems studies have contributed to generate specifications towards the sub-systems (GNC, communication, power, etc...). The spacecraft architecture has been optimized accordingly leading to a coherent design wrt all disciplines.
GNC algorithms have been implemented within a simulator for NRT (non real time) and RT (real time) tests within a laboratory environment and spacecraft performance has been successfully assessed through an end-to-end mission simulation.
Therefore, AEROFAST study confirms the feasibility of an aerocapture manoeuvre:
* Mission is fulfilled even when considering the worst cases (robust GNC strategy)
* Budgets have been established (mass, power...) coherent each others
No show stopper has been identified, demonstrating the interest of an aerocapture manoeuvre.
In addition, this collaboration of the 2.5 years between all the partners of the project has allowed to build a very good team, in a good spirit, efficient and acting in the interest of the project.
Project Context and Objectives:
1. Introduction
AEROFAST is a Mars aerocapture feasibility demonstration performed by twelve European companies leaded by AST-SAS as prime, and funded under seventh framework program of the European Commission.
This study planned over 2.5 years, started beginning of 2009 and ended in June 2011, for a total budget of about ~3M€.
2. Acronyms
A/C AeroCapture TPS Thermal Protection System
AST ASTRIUM TRL Technology Readiness Level
GNC Guidance, Navigation, Control wrt with regards to
NRT, RT Non Real Time, Real Time
3. AEROFAST: AEROcapture for Future spAce tranSporTation
Aerocapture is a new technology for Solar System exploration that uses a single pass through a planetary atmosphere to decelerate the spacecraft and achieve a targeted orbit (Fig-1).
Such manoeuvre saves a significant amount of mass with regard to a more conventional technique of insertion using propelled braking, and becomes really attractive when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions (sample return missions, in-situ missions and future manned missions which require spacecraft to enter and manoeuvre in a planet's atmosphere).
Figure-1: Aerocapture principle
Aerocapture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), guidance - navigation and control (GNC), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aerocapture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aerocapture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aerocapture technology up to TRL4 through a complete mission study of a Martian aerocapture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aerocapture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype a thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aerocapture phase recovery.
4. Statement of work
4.1 Industrial organisation
The AEROFAST consortium, lead by ASTRIUM-ST-SAS as prime, is composed of twelve European companies from eight countries:
* ASTRIUM-ST SAS (France)
* ASTRIUM-ST Gmbh (Germany)
* DEIMOS Enhgaria (Portugal)
* Amorim Cork Composites S.A. (Portugal)
* SAMTECH (Belgium)
* University of Roma - La Sapienza (Italy)
* STIL-BAS (Bulgarian Academy of Sciences) (Bulgaria)
* Institute of Aviation (Poland)
* Space Research Centre - Polish Academy of Sciences (Poland)
* ONERA (France)
* KYBERTEC (Czech Republic)
* Faculty of Sciences of the University of Lisbon (Portugal)
4.2 Work breakdown structure
The work breakdown structure is shared out of five work packages in order to fulfil the objectives of the project, and a sixth work package with deals with management, dissemination and exploitation. Details are reported Tab-1.
Table-1: WP breakdown
The work package leaders are:
WP1: AST-ST SAS WP4: DEIMOS Enhgaria
WP2: AST-ST GmbH WP5: University of Roma
WP3: Amorim Cork Composite WP6: Kybertech and AST-ST SAS
4.3 List of deliverables
The deliverables have been released according to the deliverable item list reported next Tab-2. These documents are available on the EC FP7 server.
Table-2: Deliverables items list
4.4 Planning
The entire milestone has been passed according to the schedule.
The final presentation has been done in June, 28th 2011. The last deliverables have been sent end of August 2011, as agreed by the EC. From the technical point of view the AEROFAST contract was ended at T0 + 30 months. From the contractual point of view the contract is finished at T0 + 33 months.
Figure-2: planning of AEROFAST study
Project Results:
1. Introduction
AEROFAST is a Mars aerocapture feasibility demonstration performed by twelve European companies leaded by AST-SAS as prime, and funded under seventh framework program of the European Commission.
This study planned over 2.5 years, started beginning of 2009 and ended in September 2011, for a total budget of about ~3ME.
2. Acronyms
A/C AeroCapture IMU Inertia Measurement Unit
AGL Above Ground Level L/D Lift-over-Drag
AST ASTRIUM LGA Low Gain Antenna
BC Back Composite LOS Line of Sight
CAC Cruise & Aerocapture Composite MO Mars Orbiter
DOF Degree of Freedom NRT, RT Non Real Time, Real Time
DSN Deep Space Network OBDH OnBoard Data Handling
EIP Entry Interface Point P/L PayLoads
FC Front Composite TCM Trajectory Change Manoeuvre
FPA Flight Path Angle TPS Thermal Protection System
GNC Guidance, Navigation, Control TRL Technology Readiness Level
HGA High Gain Antenna wrt with regards to
IMS Imaging Multi Spectral
3. AEROFAST: AEROcapture for Future spAce tranSporTation
Aerocapture is a new technology for Solar System exploration that uses a single pass through a planetary atmosphere to decelerate the spacecraft and achieve a targeted orbit (Fig-1).
Such manoeuvre saves a significant amount of mass with regard to a more conventional technique of insertion using propelled braking, and becomes really attractive when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions (sample return missions, in-situ missions and future manned missions which require spacecraft to enter and manoeuvre in a planet's atmosphere).
Figure-1: Aerocapture principle
Aerocapture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS), guidance - navigation and control (GNC), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements.
Currently, Technology Readiness Level (TRL) of aerocapture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aerocapture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aerocapture technology up to TRL4 through a complete mission study of a Martian aerocapture.
The objectives of AEROFAST project are:
* OBJ1: Define a project of aerocapture demonstration.
* OBJ2: Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5.
* OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms,
* OBJ4: Demonstrate/prototype a thermal protection system for such a mission
* OBJ5: Define on-board instrumentation for aerocapture phase recovery.
4. Main technical achievements
4.1 Mission overview
The frame of the AEROFAST mission has been defined as a low cost mission. The launch and transfer orbit injection is supposed to be performed by a Soyuz Fregat departing from Kourou, and able to perform the coast phase needed to get rid of declination limitations.
The complete mission sequence contains the phases of:
* Cruise & Pre-aerocapture phase: Challenge is to master the attitude/position of the spacecraft (S/C) before manoeuvre:
- Cruise : Several trajectory correction manoeuvres (TCM) are foreseen to allow Martian rendez-vous with the use of star trackers for navigation
- Approach: Before arriving to Mars, the Deep Space Network (DSN) navigation will be hybridized with an optical + radiometric navigation increasing the accuracy of targeting the aerocapture corridor.
* Aerocapture phase: Used to brake the incoming velocity thanks to the drag exerted by the atmosphere and will insert the spacecraft on the appropriate elliptical orbit.
* Post aerocapture phase: to target a quasi circular sun-synchronous orbit at a low altitude.
- Science Orbit Acquisition: Before reaching its operational orbit, the spacecraft will perform some transition manoeuvres that, with the help of the drift due to gravitational perturbations, will drive it to its final orbit.
- Science Orbit Operation: Once inserted in its final orbit, the vehicle will begin its observation mission (swath width of 25km with a resolution of ~1m) and will only perform small orbit adjustment manoeuvres.
4.2 Mission analysis
4.2.1 Mission constraints
Soyuz from Kourou offers a range of payload compatible with AEROFAST type of mission, and the flexibility to reach different declination with the re-ignition capability of Fregat upper stage.
From the mission analysis, a set of constraints reported Tab-3 has been identified which drives the cruise and pre aerocapture phase.
Constraints Justification
RQ-1 Soyuz performance must be greater than 1450 kg (1250kg spacecraft + 200kg ACU) for the launch opportunity (infinite velocity and declination at departure
RQ-2 Infinite velocity at arrival must be lower than 4 km/s To avoid energy dissipation issue during aerocapture (aerocapture design constraints)
RQ-3 The angle between the infinite velocity and the Earth direction must be lower than 90° at Mars Arrival To have aerocapture phase placed on the side facing the Earth, with the communication link available (even if Mars is hidden by the Sun at space craft arrival), and with operational transmitters during aerocapture phase (pre-aerocapture communication constraints)
RQ-4 The Sun-Earth-Probe (SEP) angle must be greater than 7° from 20 days before Mars arrival
RQ-5 The angle between the Earth direction and the vehicle velocity has to be greater than 40° during the whole aerocapture
RQ-6 The angle between the infinite velocity at arrival and the Sun direction must be greater than 45° To be able to use optical measurements (stars trackers, camera) on Mars moon in the last 20 days before re-entry (pre-aerocapture navigation constraints)
RQ-7 Pre-aerocapture navigation performance and propagation of S/C release down to EIP must be such that FPA errors at entry point are below ±1.24 deg. To allow re-entry within the aerocapture corridor (aerocapture guidance constraints)
Table-3: Constraints from mission analysis
4.2.2 Cruise & Pre-aerocapture phase
The analysis of the Earth to Mars opportunities has been performed for the whole decade 2020-2030. Considering the constraints reported Tab-3, a set of 21 opportunities has been identified (Fig-3).
The best launch opportunity from Earth has been found in Mai, 11th of 2026, with a Mars arrival expected in June, 11th of 2028 as the best compromise between minimum departure infinite velocity and transfer duration from Earth to Mars (opportunity n°9 among 21).
Figure-3: Constraint compliance map (white = no compliance)
Figure-4: AEROFAST transfer from Earth to Mars
The features of the selected trajectory are:
* A transfer of type T4 which suppose to have more than one revolution about the sun, with 6 TCM to meet the target conditions at Mars arrival (Fig-4). TCM have been optimised wrt the navigation process and the overall delta-V required to be minimized. Last TCM is foreseen not later than EIP -4h
* An injection into polar orbit at Mars by any of the two possible arrival branches (north and southbound)
* An Entry Interface Point (EIP) defined at 120km AGL, with a Flight Path Angle (FPA) of -10.74 deg.
Zone Transfer type Min. depart. velocity (km/s) Declination
(deg) Departure date Arrival date Transfer time (months)
1 T1/T2 4.3057 -10.786 31/07/2020 14/07/2021 11.43
2 T1/T2 3.7219 14.880 15/09/2022 30/09/2023 12.48
3 T1/T2 3.3342 19.735 04/10/2024 13/09/2025 11.30
4 T1/T2 3.1070 32.894 05/11/2026 13/10/2027 11.24
5 T1/T2 3.0002 28.250 10/12/2028 05/11/2029 10.84
6 T4 2.9756 29.130 05/12/2021 02/03/2024 26.87
7 T4 2.9811 19.921 04/01/2024 04/03/2026 25.95
8 T4 3.2276 10.642 29/01/2026 20/02/2028 24.71
9 T4 2.7816 -14.641 11/05/2026 11/06/2028 25.03
10 T4 3.5997 19.114 21/09/2028 27/04/2031 31.15
11 T5 3.1335 -49.399 19/04/2021 29/05/2024 37.32
15 T6 3.4743 6.046 16/02/2021 10/06/2024 39.75
16 T6 3.3289 21.477 06/10/2021 19/09/2025 47.44
17 T6 4.6517 -5.999 27/01/2023 15/04/2026 38.57
18 T6 3.2223 29.179 20/10/2023 17/09/2027 46.92
20 T6 3.1869 32.772 08/11/2025 18/09/2029 46.32
Table-4: AEROFAST selected trajectory
4.2.3 Aerocapture phase
The aerocapture feasibility is mainly driven by the conditions at the EIP (Tab-5). The aerocapture corridor is bounded by extreme over-shoot (hyperbolic exit) and under-shoot (crash) trajectories. The small width of the corridor leads to maximum FPA errors bellow ±1.24 deg at EIP.
Entry Interface Point (EIP) at 120km AGL
Relative velocity: 6377.6 m/s
Flight Path Angle: (local geodetic frame) -10.74°
Azimuth of relative velocity: -19.59°
Latitude: -77.8°
Longitude: (geodetic frame) 10,6°
Table-5: EIP features
Figure-5: Entry corridor for Aerocapture at Mars arrival
As depicted Fig-6, the aerocapture phase allows targeting the orbit inclination while exiting the atmospheric phase, but aerocapture doesn't allow meeting all in plane parameters, in particular pericenter altitude of the orbit for science operation (Tab-6).
At the end of the aerocapture at 120km altitude, spacecraft exit parameters (relative velocity ~3500m/s, FPA ~3°) lead to the nominal apocenter altitude (345km AGL) but a too low pericenter altitude (-44km ->crash).
Thus, a fast correction manoeuvre must be performed just after aerocapture to reach the nominal pericenter altitude required (290km AGL).
4.2.4 Post Aerocapture phase
The orbit for sciences operations must comply with the requirements of the scientific optical payloads. AEROFAST orbit reported Tab-6 has been chosen in order to have an almost global coverage with repeating ground track and constant lighting conditions:
* Repeating orbit condition, which allow to have image at the same location with similar light incidence condition
* Maximum surface coverage, which allow to observe the largest part of Mars
* Sun synchronous condition, which allow to compensate the Mars motion around the Sun to keep same lighting conditions
* Frozen orbit condition, which correspond to an orbit with parameters having small rates of variation in order to reduce the cost of the orbit control
Semimajor axis 3713,87 km Pericenter altitude 290.55 km
Eccentricity 7,355e-03 Apocenter altitude 345.18km
Inclination 92,7° Orbital Period (OP) 6871s = 1.91 h
Pericenter argument 270° Mars global coverage 1098 OP
Table-6: orbit features for sciences operations
4.3 Guidance, Navigation and Control architecture
The GNC architecture is composed of sensors, actuators and a guidance & control philosophy to meet the navigation requirements of each mission phase. The reference used for GNC studies is the baseline spacecraft (shape, mass) reported §6.
4.3.1 Cruise & Pre Aerocapture phase
During the cruise phase, the spacecraft is in sun pointing mode. The only changes in spacecraft orientation are for the 6 TCM and the final target orientation.
Spacecraft navigation relies on a class-3 Inertia Measurement Unit (IMU) and star trackers during cruise phase coupled with a deep space network transponder. Sun sensors may also be activated in a safe mode.
Control is done by reaction inertia wheels (for attitude control) and thrusters (for TCM and inertia wheels desaturation). For Mars final approach (EIP-5days), the optical camera is pointed towards Mars (LOS measurements) and radiometric measurements are activated to meet the requirements at EIP.
Monte Carlo simulations, including external perturbations and sensors performances, have demonstrated the robustness of the GNC architecture with accuracy < 3km (3σ) at EIP for a total delta-V < 30m/s for pre-A/C phase.
4.3.2 Aerocapture phase
Aerocapture phase must be safe (no crash or hyperbolic exit) while being complaint with the mission constraints (pre-A/C) and requirements (post-A/C). The philosophy consists in decoupling in-plane & out-of-plane motions for respectively apoapsis and inclination control, while remaining as much as possible in the aerocapture corridor (semi permeable domain).
Figure-6: In-plane & out-of-plane required motions
Figure-7: Nominal guided trajectory (not sizing)
Spacecraft navigation during aerocapture relies on the IMU (maximum deceleration about ~4g in worst case), whereas Control is performed by a set of thrusters which maximum torques have been assessed at 200N.m around the yaw axis, 100N.m around pitch and roll axis, with a minimum impulse bit (MIB) < 25ms.
The guidance performance has been assessed by Monte Carlo simulation (1000 runs, 3DOF). Results show no crash, no hyperbolic exit, an inclination well controlled and a 100% fulfilment rate of the total delta-V < 170m/s (Fig-8) for aerocapture phase and science orbit acquisition.
Figure-8: Aerocapture robustness (MtC simulations)
The performance of the control has been assessed (Fig-9) for the nominal trajectory leading to oscillations amplitude below ±5° for both angle of attack and sideslip angle (< ±1° for Pdyn >100Pa), and bank angle manoeuvres performed with accuracy better than 5°.
Figure-9: Control performance during nominal guided trajectory
4.3.3 Post Aerocapture phase
Just after aerocapture, TPS heat shield has to be jettisoned to allow fast correction manoeuvres to target the in-plane parameters of the science operation orbit:
* During the science orbit acquisition, navigation relies on IMU. The two burns which are required at apocenter & pericenter (mainly to shift pericenter altitude from -44km to 290km) are performed by the attitude control thrusters of the Mars Orbiter.
* Once Mars Orbiter placed on the science operation orbit, navigation relies on the star trackers and the IMU. The attitude control is performed by reaction inertia wheels, while the slow decrease of the semi-major axis (Fig-10) is corrected by bi-impulsive corrections done by the thrusters every 61 orbits to limit ground track drift at 30m maximum at the equator wrt the nominal ground track.
The science orbit control requires a delta-V~3.6m/s for 1 period of repetition (1098 nodal periods) corresponding to a full coverage of the Mars surface.
Figure-10: Evolution of semi major axis (if no correction) compared with the stability of the argument of perigee over an entire period of repetition of 1098 nodal orbital periods (frozen orbit conditions)
4.4 Spacecraft design
Missions with aerocapture phase generally involve spacecraft architecture based on two main modules, an Orbiter module encapsulated within a protective module jettisoned after Aerocapture.
Figure-11: Product tree
AEROFAST spacecraft product tree depicted Fig-11 is based on three different modules, a Mars Orbiter, a Cruise and Aerocapture Composite composed of a Front Composite and a Back Composite (option):
* During pre-aerocapture and aerocapture phase, the Mars Orbiter module is implemented within the Cruise and Aerocapture Composite (Fig-12)
* After aerocapture, prior to post-aerocapture manoeuvres, Front & Back Composites are jettisoned (Fig-12)
* Post-aerocapture manoeuvres for sciences orbit acquisition & operations are performed by the Mars Orbiter (Fig-12)
Figure-12: Aerofast different configurations
4.4.1 Aero-shape trade-off
Aerocapture requires an aeroshape that provides a lift-over-drag (L/D) ratio with sufficient provision (L/D>0.3) and sufficient static stability performances. In addition, the aeroshape must protect the payloads (P/L) from the severe aerothermal environment during the aerocapture, and shall be safely jettison able at the beginning of the post-A/C phase.
In the frame of AEROFAST, a trade-off has been performed considering two configurations:
* A blunt body capsule derived from Apollo, ARD concepts
* A lifting body dealing with bi-conic shape
By considering L/D requirements, payload protection wrt aerothermal heating and layout constraints, the bi-conic shape has been selected.
4.4.2 Aerothermodynamics analysis of the biconic shape
Aerodynamic performances of the bi-conic shape have been assessed for different angle of attack by Euler calculations, results exhibiting L/D ratio > 0.4 at angle of attack < 40deg (Fig-13).
Figure-13: Max dynamic pressure point - Mach 27
Figure-14: Satellite flow interaction t=91s, 40° angle
Spacecraft-flow interactions (Fig-14) have been analysed, demonstrating (Fig-15) the need of a Back Composite protection for the Orbiter (locally heat fluxes >20kW/m2 mainly due to flow radiation)
Figure-15: Radiation heat flux on Orbiter at t=91s
Figure-16: Max heat flux history (T°w=300K)
Heat fluxes histories (Fig-15) have been assessed along the two worst case aerocapture trajectories identified by Monte Carlo analysis:
* overshoot trajectory corresponding to the maximum heat flux trajectory, with a maximum flux ~1064KW/m2 found at stagnation point (turbulent regime applied over the whole surface wrt Reynolds criteria)
* undershoot trajectory corresponding to the maximum energy trajectory
The baseline TPS foreseen to protect the spacecraft is the Norcoat liège material (used on ARD, Beagles), glued on Front & Back Composite structures. The TPS thickness is driven by maximum temperature at the interface with the structure (T°max ~450K driven by the silicone based glue).
3D computations results considering Norcoat pyrolysis/charring and ablation lead to:
* thickness from 19mm to 6mm over the Front Composite surface, with a maximum ablation of 10mm
* thickness of 3.3mm uniform over the Back Composite surface
4.4.3 Spacecraft architecture
Spacecraft architecture results from trades-off performed at system level, with permanent drivers for saving mass and increasing robustness by selecting principles as simple as possible.
A trade-off has been conducted about the Cruise Stage. Several options were assessed: 2 configurations with external Cruise Stages jettisoned before aerocapture and 1 configuration with an integrated Cruise Stage. At last, the integrated Cruise Stage has been selected has the best compromise between mass saving and robustness (no additional separation mechanism).
The baseline architecture of the spacecraft (CAC and MO) is depicted Fig-17.
Figure-17: Baseline architecture of AEROFAST CAC and MO
The design and the layout of the spacecraft are driven by the mission sequence:
* Spacecraft must comply with the launch vehicle fairing and interplanetary orbit injection mass capability (Soyuz from Kourou)
* Power generation & storage, communication link to Earth must be available during cruise phase and science orbit acquisition & operation phases
* Autonomous guidance, navigation and control must be available along the entire mission from cruise, approach, aerocapture, post aerocapture phases.
A preliminary thermal control analysis has been performed which concluded that no active thermal control system is required for on board equipments. However, a cold case has been identified on hydrazine tanks which may require active heater during cruise.
The functional architecture of the spacecraft depicted Fig-19 has been established with the desire to make re-use of components as much as possible, and to fulfil redundancy requirements.
4.4.4 Spacecraft layout
Most parts of the power, On-Board Data handling (OBDH) and communication systems are integrated within the MO. The equipments which are solely needed during cruise and aerocapture phases are installed within the CAC. The equipments required for the entire mission which need to have outside access have been duplicated and implemented both within CAC and MO (solars array, communication antenna and attitude thrusters).
The specifications of the onboard equipments (measurement accuracy, torques, power budget etc...) have been established wrt the system studies, and by considering the worst cases identified by the Monte Carlo analyses.
Attitude & orbit control subsystem
Navigation:
* Star trackers and sun sensors both on CAC and MO
* IMU (NG LS-200S) for aerocapture and orbit acquisition within MO
* Navigation camera for approach phase on CAC
Control:
* Reaction wheels for cruise and orbital phase within MO
* Thrusters:
- 6 aerojets MR120 (90N) on CAC for aerocapture + 4 CHT-5 (5N) for reaction wheels desaturation during cruise
- 12 CHT-20 (20N) on MO for science orbit acquisition, attitude control, and inertia wheels desaturation
* 1 tank of Hydrazine (N2H4) is required within CAC to cover cruise, pre-A/C and A/C needs (113kg propellant), whereas 2 tanks are required within MO to cover post-A/C needs (~28kg)
Figure-18: Equipments considered
Figure-19: AEROFAST functional architecture
Power subsystem
The power architecture is based on a hybrid approach, a Direct-Energy-Transfer (DET) for cruise phase which is quasi permanently sun illuminated and a Peak-Power-Tracking (PPT) for orbital phase which is dominated by alternative eclipses/daylights in short period:
* During cruise & pre-A/C phases, power needs are in average ~115W which are provided by solar cells of ~1,23m2 fixed on the CAC surface.
* For orbital phase, power needs vary from 143 watts (eclipse) to 562W (needs for observation and transmission during daylight) provided by 2 solar arrays of 5,6m2 each on MO.
Batteries (8x VES 140 cells) are implemented within the MO with a total capacity of 39A/h at 28,8V.
Figure-20: Batteries and capacities
Communication subsystem
The communication subsystem provides a data link for housekeeping, telemetry/telecommand and transmission of scientific data.
* 4 Low Gain Antennas (LGA) are used for telemetry/telecommand during cruise and for orbit acquisition phases. LGA are implemented both within CAC and MO, and in a redundant way (mission critical).
* 1 High Gain Antenna (HGA) is used for transmitting scientific data to Earth. Antenna diameter is foreseen to be 2.1m diameter, with a transponder operating in Ka-band at minimum data rate of 594kbit/s. HGA is implemented within MO.
Scientific payload
Science payload is assumed to be 125kg maximum, with a maximum power consummation of 100W.
For the moment, an Imaging Multi Spectral (IMS) sensor is foreseen (87,5kg, 55Watts) dedicated to take pictures for a global observation of Mars surface (panchromatic scanner: 500-800 nm; multispectral scanner: 3 channels: Green, Red, NIR; Spatial resolution, IFOV: 1m (Pan); 4m (Multi) in NADIR position; swath ~23km)
Instrumentation is also foreseen for environment characterisation and TPS analysis during aerocapture phase (see §5.6).
4.4.5 Mass breakdown and geometry
Mass breakdown, geometry and centring are depicted Fig-21 and Tab-7.
Table-7: AEROFAST mass breakdown
Figure-21: CAC shape and centring box
4.5 Validation of the GNC performances
4.5.1 Methodology
Based on the mission, architecture and general design of the spacecraft, the GNC algorithms performance for Pre-A/C and A/C phases has been assessed through 6DOF simulators with Non Real Time (NRT) and Real Time (RT) breadboards.
Both pre aerocapture and aerocapture performances have been assessed separately. The logic is depicted next Fig-22.
Figure-22: development flow
As explained previous sections, GNC architecture relies on both optical (1 camera) and radiometric measurement activated 5 days before EIP, last correction manoeuvres being feasible until 4h before EIP.
4.5.1 Pre aerocapture GNC assessment
4.5.1.1 Physical mock up (IP-lab)
Optical navigation is a subsystem to GNC that can provide information on spacecraft (S/C) location and attitude with respect to a celestial body. In AEROFAST context, this applies to the Pre-Aerocapture mission phase. Main optical observables are the Line-of-Sight (LOS) to a target and image dimensions (diameters for spherical bodies).
To verify the performance of such observables it is necessary to recreate (within a valuable scope) the imaging conditions. Therefore, the IP-lab mock-up depicted Fig-23 has for objective to simulate the mission profile with real images in order to test imaging and GNC combined performances.
Images are generated by a live camera looking at a ball lit by a source with adequate irradiance. The produced image does simulate the solar illumination and the range of Mars planet.
Figure-23: IP lab (Mars/Sun unit)
The calibration and test campaign has allowed validating the IP algorithms in relevant cases with real camera performances.
4.5.1.2 Pre aerocapture Non Real-Time GNC performance (FES)
NRT Monte Carlo simulations (200 runs) for Pre-A/C phase include sensibilities to initial position and velocity errors at EIP-5days, as well as to mass, propulsion and relative biases between camera and attitude sensors.
Results reported Tab-8 show that conditions at EIP are met even by considering worst case (initial position and velocity errors, increased misalignments) which validates pre-A/C GNC architecture. The simulations indicate that relative bias is the most important contributor to the GNC performance. A relative bias of 72µrad (3σ) leads to comply with the requirement at EIP of 3km (3s), a contrary to a bias of 290µrad (3σ).
Figure-24: Pre Aerocapture NRT simulation (FES) and test cases selected
Figure-25: Spacecraft position dispersion wrt target position at EIP (XY and XZ view)
(MC2 test case)
Figure-26: Statistic on pre-A/C position and velocity errors, FPA dispersions wrt target FPA (MC3 test case)
Table-8: Pre A/C NRT tests results
4.5.1.3 Pre aerocapture Real-Time GNC performance (HIL)
The Real Time simulation is performed to prove the GNC capability to run on a LEON3 processor in a nominal way. The simulation starts 5 days before EIP. Two cases have been selected from the NRT Monte Carlo simulations: Nominal and Robustness case.
Figure-27: Pre A/C RT tests results (Robustness case)
Requirements are fulfilled and results are coherent of the NRT simulations for both cases, demonstrating the feasibility & robustness of the autonomous pre A/C GNC concept.
4.5.2 Aerocapture GNC assessment
GNC design for aerocapture is based on:
* For Guidance: a numerical apoapsis predictor-corrector derived from MSR-O and ATPE studies
* For Navigation: A5-like inertial navigation with autonomous alignment (class-3 IMU)
* For Control: a flight proven ARD-like algorithm
4.5.2.1 Aerocapture Non Real-Time GNC performance (FES)
NRT Monte Carlo simulations (6DOF, 500 runs each) have been performed for A/C phase in order to assess GNC performance and robustness by considering a set of dispersion (EIP, atmosphere, aerodynamic coefficients, MCI, IMU).
Results reported Tab-9 shows 3 violations of the requirements.
* Dynamic pressure (~6.4Kpa instead of 6Kpa max)
* AoA at low pressure (5.4° instead of 5° max)
* AoA at high pressure (1.8° instead of 1° max)
But these violations which occur for some dispersed cases are very slight and not considered as critical.
Table-9: A/C NRT tests results (Monte Carlo 6DOF, 500runs)
4.5.2.2 Aerocapture Real-Time GNC performance (PIL)
The Real Time simulation is performed to prove the GNC capability to run on a LEON3 processor in a nominal way as for Pre-A/C phase, and to check the GNC needs wrt the available resources (CPU, delays, memory) as dynamic phenomenon are highest during the A/C compared to the pre-A/C phase.
The RT results reported Tab-10 are in line with the NRT A/C simulation results.
Table-10: A/C RT tests results (nominal)
GNC Real Time results are good even if simulations have revealed impact of the GNC computational delays, mainly guidance computation time (up to 0.8 s). Maximum CPU load vary from 65 % (nominal) to 90 % (worst-case)
4.6 Instrumentation
A preliminary instrumentation plan has been established in order to cover the need of data for three different objectives: environment analysis, TPS analysis and scientific payloads.
For each proposed sensors, compatibility with the environment and performance wrt the requirement has been assessed. In addition, a functional architecture has been proposed based on a sensor network built on two levels: measuring cells, containing different types of sensors and cluster controller, collecting measurements and communicating with the on board computer.
4.6.1 Environment analysis sensors
The sensors are dedicated to provide data about the environment during the re-entry phase. The two categories of sensors foreseen are integrated pressure and heat flux sensor (Raflex type) and temperature sensors implemented within thermo plugs.
The Raflex type sensor is used to measure both pressure and heat flux. It is composed of a pressure tap and a calorimeter made of Molybdenum as depicted Fig-28a. Thermo plugs are made of 3 thermocouples each placed at different depths (Fig-28b).
Figure-28 : RAFLEX sensor (a), thermo plug (b)
5 Raflex sensors and 16 thermo plugs are foreseen on the front composite.
4.6.2 TPS analysis sensors
The main interest consists in collecting information on the plasma state and the chemical composition of the TPS surface during the aerocapture phase. Spectrometric measurements are foreseen by using pyro-sensors & several filters as depicted Fig-29.
Figure-29: Pyro sensor
5 pyro sensors are foreseen according to 5 different generatrices.
4.6.3 Scientific payloads
The goal of the mission after the aerocapture phase is the global observation of the Martian surface by pictures taken by an optical payload and transmitted to Earth.
The selected payload is an Imaging Multispectral Sensor (IMS) whose design and performances are depicted Fig-30.
Figure-30 : Scientific payload (IMS)
4.7 TPS improvments
Norcoat liege TPS (cork based) has been chosen as the baseline for Aerofast mission wrt the demonstrated performances of this TPS for atmospheric entry probes (ARD, Beagle-2).
Nevertheless, research and development of advanced cork based TPS have been initiated in order to improve mass, performance, resistance to erosion, resistance of the char material and surface aspect after degradation. The objective is to propose reinforced material for front shield and superlight material for back shield. Several tracks have been followed including:
* different cork granules
* different types of resins (phenolic, furan)
* different types of fillers (carbon fibres, basalt fibres, glass spheres)
* optimisation of mixing process
Preliminary basic characterisations leaded to select the most promising formulations for complementary plasma tests in order to assess thermal performance and ablation, for comparison with Norcoat liege & P50 reference materials.
Two mission profiles have been selected corresponding to the maximum energy and heat fluxes trajectories. Samples have been manufactured for plasma tests (Comete facility).
Figure-31: selected material for plasma tests campaign
Globally, behaviour of all material is quite good, with better surface aspect for both mission profiles. Masse losses are of same order of magnitude, but for few samples, thermal insulations seems not so good (those with carbon fibres). Best results have been found with the TPS3J formulation (phenolic + carbon fibres)
Figure-32: Plasma tests on Cork based material
Therefore, tests demonstrated that reinforcement with fibres improves the resistance to ablation, confirmed that phenolic grade resin is still the most appropriate compared with new resins tested (green resins).
The manufacturing of a very complex prototype (small radius, thickness not constant as depicted Fig-33) demonstrated that cork based TPS offer wide possibilities for moulding or machining panels, but expertises have shown important density heterogeneities. Therefore, complex shapes are achievable with cork based TPS, but manufacturing process has to be optimised accordingly.
Figure-33: TPS prototype
A non ablative option has been also assessed. A preliminary design based on C/SiC TPS for FC and FEI TPS for BC leads to a total TPS mass of about ~150kg, which is similar to the TPS mass budget with Norcoat liege. But, heat-shield robustness and maturity based on C/SiC TPS is lower compare with cork based TPS. C/SiC could be interesting only if FC underlying structure could be removed (mass saving ~75kg).
5. Assessment of the objectives
AEROFAST project has been established wrt five top objectives
* OBJ-1: Define a project of aerocapture demonstration -> completed
- Mission analysis has been performed -> set of requirement for system studies (ATD, Aero, GNC) -> set of requirements for sub systems (spacecraft design, propulsion, power etc...)
- A coherent project of an aerocapture has been defined (mission, system, sub-system)
* OBJ-2: Increase TRL up to 4/5 of the planetary relative navigation and the aerocapture algorithm -> completed
- TRL5: breadboard validation within a representative environment.
- Algorithms for pre-aerocapture and aerocapture phases have been developed according to the spacecraft design, and performance has been assessed through breadboard simulations (NRT, RT)
* OBJ-3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC algorithms -> partially completed
- Objective closely linked to the objective-2. Simulations in RT of the GNC algorithms and the hardware (IP-lab) have been successfully validated but separately.
* OBJ-4: Demonstrate/prototype the TPS for the mission -> completed
- Usual cork based TPS is suitable for an aerocapture mission
- Non ablative TPS comparable in term of mass to the ablative material
- Improvement of cork based material has been performed (roughness improvement of the charred surface) with the addition of fibres
* OBJ-5: Define O/B instrumentation for aerocapture phase recovery -> completed
- On-board instrumentation has been defined for: navigation, TPS/environment characterisation and scientific observations
6. Conclusions
AEROFAST is not dedicated to perform a phase-A, but to prepare for such a mission demonstration by increasing technology maturity to TRL4 (mainly GNC & thermal aspects).
After two years of studies, the overall mission has been defined and the conditions (EIP...) required to perform each phase (cruise, pre-A/C, A/C, post-A/C) established. Mission and systems studies have contributed to generate specifications towards the sub-systems (GNC, communication, power, etc...). The spacecraft architecture has been optimized accordingly leading to a coherent design wrt all disciplines.
GNC algorithms have been implemented within a simulator for NRT and RT tests within a laboratory environment and spacecraft performance has been successfully assessed through an end-to-end mission simulation.
Therefore, AEROFAST study confirms the feasibility of an aerocapture manoeuvre:
* Mission is fulfilled even when considering the worst cases (robust GNC strategy)
* Budgets have been established (mass, power...)
No show stopper has been identified, demonstrating the interest of an aerocapture manœuvre.
In addition, this collaboration of the 2.5 years between all the partners of the project has allowed to build a very good team, in a good spirit, efficient and acting in the interest of the project.
Potential Impact:
From the past experience on aerocapture, this transportation technology is clearly identified as core technology for the upcoming future missions and will save significant mass at launch. The past activities highlighted the issues and the complexity of such a transportation technology. TRL is in the order of 2 to 3 in Europe. In order to use the aerocapture technology for operational missions while mitigating future development risks, the TRL of such a technology must reach the level 6. TRL6 shall be reached through a flight demonstration.
Such a demonstration can only be envisaged at European level and shall be led by ESA, the European Space Agency. Assuming that the overall development of such a mission is in the range 400 to 600 M€, 5% of the budget is necessary for the feasibility assessment (called phase A). The approach taken by EC is to help Agencies to prepare for the future and to fund space technologies supported by Agencies with clear application on future programmes and those Agencies cannot fund (too long term, too risky ...).
So clearly AEROFAST is in line with this approach: AEROFAST end goal is not to perform a phase A of the mission demonstration but is to prepare for such a mission demonstration and increase the TRL level: AEROFAST intends in the frame of the FP7 to reach a TRL 3 to 4 for this technology. Moreover the AEROFAST partners have been chosen in accordance with these goals. The AEROFAST team is not tailored to run a phase A of a mission demonstration. Anyway from the cost point of view the budget dedicated to AEROFAST does not allow building and performing an overall feasibility assessment of the demonstration. It might be précised that the consortium worked very well together and is spread over eight countries. At least these different partners from these different countries would be really enthousiastic to go further on the project and develop it together until flight.
Following tasks were envisaged within AEROFAST to increase the TRL level up to 3 / 4:
First of all a mission were drafted and mission and design requirements were assessed. This provided the framework of AEROFAST and the conditions for the other tasks.
Secondly, AEROFAST worked in parallel on several topics necessary for TRL increase:
* Increase our knowledge on aerothermodynamics and flow behaviour,
* Improve and better assess the effect of ablation with respect to the centre of gravity,
* Make a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5,
* Demonstrate/prototype the thermal protection system for such a mission.
In addition it might also be precised that the different areas of research of AEROFAST could also be more widely used and results obtained used for other applications. Table n°1 highlights the growth potential of AEROFAST results to other applications and identifies the additional tasks to be implemented. AEROFAST is built and studied for a demonstration in a Martian environment, nevertheless interest is to identify whether some of the disciplines could be adapted to other applications:
* Disciplines linked to the necessity of having an atmosphere, and possibly different from Mars,
* Generic disciplines.
Growth potential to
Disciplines & products: Mars application Another atmosphere generic tasks
Mission
vehicle design:
heat shield
cruise module
orbital module X
X
Valid for the cruise module and for the orbital module - to be adapted to the mission needs
aerocapture corridor X X Methods to compute corridor can be re-used
Corridor in Earth different from the one on Mars
Aerodynamics X CO2 environment with winds & storms on Mars
Very different physical phenomena
Shape of the front shield might be different from one atmosphere to another
Aerothermal Environments X
Innovative thermal Protection System X X Material based on cork does not fit for an aerocapture on earth (due to very high thermal environment)
Results obtained could be assessed in other environmental conditions in order to test their behaviour, for instance an re-entry from the ISS orbit (max 400 km)
Necessity to perform computations and associated ground tests in the facilities
Aerocapture GN&C systems
associated simulator
test bench X X for G
X Guidance algorithms independent from the atmosphere
navigation and control choices pending upon Mars
test bench generic, only sensors to be adapted to the mission to be tested
Disciplines & products: Mars application Another atmosphere generic tasks
On-board navigation systems X X Optical navigation to be adapted to Earth because of the planets sizes
System could be re-used for another planet, no interest around earth (GPS)
Innovative on-board instrumentation X Instrumentation for the aerocapture phase is dedicated to Mars
Table n°1 Utilisation of AEROFAST results to other applications
More globally speaking AEROFAST participates:
* to sustain a competitive industry (including manufacturers, services providers) in defining an aerocapture demonstration on Mars while making a significant progress in space transportation by increasing the TRL of the planetary relative navigation and the aerocapture algorithm up to 5, as well as prototyping an innovative thermal protection system for such a mission and innovative on-board instrumentation for aerocapture knowledge improvement,
* To provide appropriate services and infrastructures.
Moreover the project AEROFAST fully agrees
* as well with a progress towards the sustainable provision of technologies needed by the European Space to become non dependent,
* As with the consolidation of long term sustainability,
* as with the improvement of the economical aspects of a domain known to be demanding in terms of reliability, experimenting novel techniques (relative navigation) and methodologies (thermal protection system).
Regarding Space independence, this aerocapture technology will be a potential solution for several future applications:
* Sample return missions,
* Manned missions returning from the Moon,
* Robotic missions towards planets such as Mars, Venus or Titan.
Those potential applications are inline with the European space policy in terms of space transportation and ESA approach regarding future space exploration: through Moon and Mars Missions.
All countries (USA, Europe and Russia) agree on the importance and interest of aerocapture for future high mass cargo missions and manned missions, to allow for human expansion into the solar system. All past studies and projects show that such transportation vehicles must rely on aerocapture to be mass effective: using atmospheric drag to slow transportation vehicles is regarded as one of the largest contributors to making both lunar and Martian missions affordable.
The MSR mission and future manned missions are scheduled not earlier than 2022, so roughly speaking in more than 15 years. On Moon side some early robotic missions are foreseen in order to improve the necessary technologies for MSR, focused on rendezvous and capture and soft-landing - aerocapture here is not addressed. NASA would also envisage manned missions on the Moon around 2020 where aerocapture would become of interest for returning the crew to Earth. Globally speaking Aerocapture transportation technology should be mastered within the next decade.
Thanks to AEROFAST a good step is likely to be achieved and allows for further improvements and knowledge on this technology.
