DUMBO - DUal Moon Beam Orbiter
Reviewed by: David
Review date: 22/07/2022
Review status: ✅DONE
As part of the 2022 ESA Concurrent Engineering Challenge (CEC), a team of students at EPFL, together with another two teams based at ESA Education Training Centre (ESEC-Galaxia) in Belgium and at the University of Luxembourg, conducted a week-long study to design a remote prospecting mission in lunar orbit.
For a list of terms and acronyms often used, check the Glossary section.
This study was performed at eSpace Concurrent Design Facility by the following team:
|Denis Galagan||eSpace||Team Leader/Facilitator|
|Mathieu Udriot||eSpace||Team Leader/Facilitator|
|Thomas Manteaux||EPFL||Attitude & Orbit Control (AOC)|
|Quenting Delfosse||EPFL||Communications & Data Handling (CDH)|
|Louise Cayroche||EPFL||Structures & Mechanisms|
|Kevin Marangi||EPFL||Structures & Mechanisms|
|L'Emira Emma Chehab||EPFL||Trajectory Analysis|
Under the responsibility of:
S. Hamel, eSpace, Study Coordinator
D. Rodríguez, eSpace, Study Coordinator
With the technical support of:
N. Cardines, eSpace, Systems Engineering Support
M. Juillard, eSpace, Systems Engineering Support
H. Koizumi, University of Tokyo, Systems Engineering Support
C. Norhadian, eSpace, Logistics & Admin Support
From April 4 - 8, 2022, an interdisciplinary group of EPFL students took part in the annual Concurrent Engineering Challenge (CEC). The CEC is a week-long event organized by the European Space Agency’s (ESA) Education Office and Systems and Concurrent Engineering Section intended to teach graduate students the principles of concurrent design for space missions and to support the development of new concurrent design facilities throughout Europe.
For eSpace, the CEC'22 was a vehicle to re-inaugurate activities associated with concurrent design at EPFL as well as an opportunity to revise the present use and configuration of the center's forsaken concurrent design facility (CDF).
The objective of this study is to design a mission involving a mothercraft and two cubesats to remotely prospect the poles of the Moon in search of water ice using bi-static reflectometry.
The mission is envisioned as a distributed space system in lunar orbit consisting of a single, larger satellite (aka the "mothercraft") and two cubesats. The mothercraft would be used as a signal transmitter of the bi-static reflectometry system; while the two cubesats would work as signal receivers. This mission concept provides increased operational flexibility and redundancy. Reflectometry provides advantages through high sensitivity, low-gain (high field of view), and moderate date volume.
As part of this study, only the mothercraft shall be designed. The two cubesats must be simply used as inputs to the study and shall be considered as dummy payloads.
As part of ESA CEC 2022, three student teams supported by experts and located at ESEC, the University of Luxembourg, and EPFL took on the challenge to design, in less than a week, a lunar remote sensing mission to characterize water ice in the polar regions of the Moon via the use of bi-static reflectometry.
Water is one of the most sought-after resources in space. Due to its multiple uses, from the production of propellant and breathable oxygen for astronauts to drinking water and radiation shielding of high energy particles, water is driving the selection of prospective areas of exploration for governmental agencies and private corporations alike in the upcoming years. Water ice also holds unique scientific value providing information on the history and chemistry of the inner Solar System (trapped Solar wind particles).
Since the 1960s water in the form of ice has been suspected to be widely located in lower temperature regions in the lunar poles. Due to the small inclination of the Moon’s equator to the ecliptic (1.5°), these regions are barely lit by sunlight or even remain permanently in shadow, acting as “cold traps” for volatile compounds. As a result of the findings made by the LCROSS and LRO missions, the lunar South Pole is being considered a prominent destination for future exploration missions. The objective: to validate in-situ de existence, amount, and distribution of water ice in the polar regions of the Moon.
In this context, the mission primary and secondary objectives are:
Similar to how on Earth oceanographic and cryospheric properties can be measured from orbit by means of GPS reflected signals, bi-static reflectometry can be used to detect, quantify, and map water ice deposits on the polar regions of the Moon.
Bi-static reflectometry takes advantage of the signals generated by the mothercraft ("the transmitter") whose reflection on the lunar surface are then measured by the the two cubesats ("the receivers"). The time delay and frequency change between the original signals from the mothercraft and the receiving signals by each of the cubesats can provide valuable information on the state and composition of the lunar surface.
The mission requirements and main system requirements, as provided by the engineering team at ESA prior to the study, are presented in the following table.
|MIS-M-001||The mission shall use Ariane 6 or Vega-C|
|MIS-M-002||The mission shall measure water ice quantity at the Lunar Polar regions|
|MIS-M-003||The science phase shall be 1 year|
|MIS-M-004||The mission shall be launched before 2027|
|MIS-M-005||The mothership shall carry and deploy 2 CubeSats in low Lunar orbits|
|MIS-M-006||The mission shall use a mothership for data processing and as communication relay to Earth|
|MIS-M-007||The mission shall use bi-static reflectometry with: a) a signal transmitter on the mothercraft and b) a receive on each of the cubesats|
Mothercraft design drivers
3U CubeSats design drivers
(out of scope for the study )
A summary of the outcome of the study is presented in this section.
The following figure illustrates the overal concept of operations (CONOPS) for the mission, dubbed DUMBO, or DUal Moon Beam Orbiter.
A Vega C launcher brings the mothercraft into a 5-day trajectory to lunar orbit. After a lunar orbit insertion burn, the mothercraft is placed in low lunar orbit, 100km above ground where the two cubesats are deployed. The total delta-v required for these maneuvers is 1.04 without taking into account the trans-lunar insertion (TLI) performed by the kickstage. Total propellant mass without TLI yields 172.2 kg.
The mothercraft has been designed with two main configurations: folded and deployed. Overall mass of the mothercraft is 1323 kg (including kickstage and adapter). Power is provided via steerable solar panels providing 800 W during sunlight. During eclipses, 590 W·h batteries provide more than 355 W and can be also used for short periods of time in other modes of operation. Internal structure is made of 3-mm thick aluminum honeycomb sandwich panels. Two standard cubesats deployment mechanisms are used for deployment of the payloads. The spacecraft is propelled by 4 Bradford ECAPS 200 N HPGP thrusters and its attitude is controlled via 8 Bradford ECAPS 0.5 N HPGP thrusters and 4 3-axis 4Nms reaction wheels. All thrusters make use of the same LMP-103S propellant. For telemetry and telecommands, a fixed X-band atenna dish is used. S-band and P-band patch antennas are used for the transmission of the bi-static signals.
Further details on the mission and the design of each subsystem are presented in the following results section.
From the high level requirements described above, each subsystem translated some into lower level, more precis requirements applicable to their design and operation. The subsystem requirements are listed at first and the final solutions, retained after several iterative loops, are explained.
The final trajectory is based on the performance of the Vega C launcher. Vega C is able to insert the mothercraft and its payload on an elliptical orbit (5700kmx250km) around the Earth. By performing a trans-lunar insertion (TLI) burn and then a lunar orbit insertion (LOI) burn with its kick-stage, the spacecraft is able to raise its apogee to 400'000km and then circularize around the Moon in low lunar orbit (LLO). The duration of the trip to the Moon's orbit is around 5 days.
Around the Moon, at 100km in low lunar orbit, the orbit duration is 7065 sec (~2h). The cubsats will be injected on the same orbit and the mother craft will manoeuvre to fall back in terms of true anomaly to have an good angle to perform the measurements.
At end of mission, when no more fuel is available for station-keeping or another constraint prevents the science objectives (that shall be at least 1 year after commissioning), the spacecraft is deorbited and crashes on lunar soil.
The total Detla v budget is shown in the table below. A contengincy on the propellant mass will be added at system level in the system budget section.
|Manoeuvres (burns)||Delva v [km/s]||Propellant mass [kg]|
|TLI (performed by kick-stage)||1.866||kick stage|
|Total (w/o TLI)||0.993||148.2|
|Sub-System contingency||5 %||-|
|Total (w/o TLI) with contingency||1.04||172.2|
Starting with the folded configuration (while in the fairing and before deployment and commissioning). The spacecraft, its large dish antenna, solar panels and other equipments are on top of the kick-stage (see structures and mechanisms). For this configuration, it is important to have a compact, inactive spacecraft, with a balanced weight distribution.
To be able to deploy the solar panels, the deployment mechanisms shall be clear of any interference.
The main structure is a cubic box 0.7 m wide. In the deployed configuration, the CoG is not exactly at the center of the box, it is offset by about 100 mm on two axes and 20 mm in the remaining axis.
At the center of gravity, the moments of inertia are 22, 74, and 79 kg·m2 in the x, y and z axis, respectively.
A 2D diagram showing approximately the distribution of the subsystems inside the deployed configuration is shown below.
The external structure (the cubic box described above) is made of aluminium, 3 mm thick, to stand the loads during launch. This makes up a mass of 16.4 kg. It is suggested to use sandwich panels (aluminium skin on the side of a honeycomb structure) for the internal structure, holding the equipment. This material is light and yet very resistant.
Several machanisms have been identified: the solar panels shall be deployed and orientable. To accommodate the addition of the kick-stage and get rid of it after its burn, a separation mechanism connects it with the main spacecraft structure as shown in the figure below. Two standard cubesat deployment mechanisms are also needed to inject the payloads.
The x-band dish antenna is fixed. The s and p-band antennas are patch so they cannot be moved neither.
For simplicity, it was decided to use the same propellant for the main propuslion system than for the AOCS. The choice was made to go for a "green" propellant, LMP-103S. This chemical monopropellant was chosen instead of hydrazine for several reasons. Hydrazine might be banned in the coming years because of its toxicity. LMP-103S has two advantages: 1) it has a higher Isp and 2) it is more dense. On the flip side, LMP-103S is so far the more expensive alternative due to lack of demand for mass-production.
The kick-stage to perform the TLI is a STAR 30 BP.
The main propulsion module is made of 4 Bradford ECAPS's 200N HPGP thrusters. 8 Bradford ECAPS's 0.5N HPGP thrusters are used for the AOCS (see below).
Other elements of the propulsion system have been identified but not precisely selected like: pressurant tank, pressure transducers and regulator, fill, drain, and Latch valves, pyrovalves, filters.
The total mass of the propulsion subsystem is estimated at 30.4 [kg].
|Reaction wheels||4||3-axis stabilization||Life of sensors, wheel bearings||4 Nms||Hot||Ref|
|Thrusters||8||Orbit manoeuvres, momentum damping||Propellant mass||0.5N||Cold||Ref|
The same table is provided with the sensors below:
|Sun sensors||6||Initial vehicle attitude, coarse attitude data||Power, radiation||+/- 0.25°||Cold||Ref|
|Star Trackers||2||Absolute altitude||Power, radiation||+/- 0.002°||Hot||Ref|
|Gyroscopes||4||Rotational rates||Power, radiation||+/- 0.3°/h drift||Hot||Ref|
The hardware selected for the subsystem is presented below.
The margins are shown to be sufficient in the figures below.
All values in the budget table below are given in Watts [W].
Other elements estimation (based on VPCDU-1 device)
Power system total mass is 60 kg. Solar arrays provide up to 800 W during Sunlight (equipment + batteries). Batteries provide more than 355 W during eclipse and can be used for short periods of time in other modes.
Insulation layer between the thrusters and the spacecraft:
Active thermal system for the solar panels:
Example of a Kapton adhesive heating patch circuit below.
Active thermal system for the eclipse mode:
Total weight of the active thermal system is around 2.1 kg.
The contingencies droped during the week as the design got more and more precise. Especially the one for structure which went from 100% to 20% thanks to more rigorous calculations of their mass.
|Subsystem||Contingency [%]||Mass w/ contingency [kg]|
|Structure & mechanisms||20||175.28|
An addition contingency of 20% is added at system level because it is still an early study. This brings the total mass to 453,29 kg. From the propulsion, one knows the propellant mass is 172.2. 20% margin is also added there to account for spin and despin of the spacecraft before and after the use of the kick stage, to get 206.64 kg.
|System||Mass w/ contingency [kg]|
|VAMPIRE 937 Adapter||120|
For a total launch mass of 1323 kg, this compared to the Vega-C performance of 1700 kg to an orbit of 5700x250 km, i=6°. This means our system can be launched by Vega-C to the required orbit!
For general lessons learned about concurrent engineering, go to the useful_tips page (accessible to eSpace staff only).