MAGIC

https://docs.google.com/document/d/1FzKE3pckUnpLt-9pGIOKBbe2K2rwpCml/edit?usp=sharing&ouid=102476260250496758607&rtpof=true&sd=true

https://drive.google.com/file/d/1fvOCksE2abGp1cWLJynF1R74y_pPB3sh/view?usp=sharing

The problem

Hi, we’re the “MAGIC” team. We chose to tackle one of the biggest problems of Venus exploration: power storage and production. Rovers are neat machines that provide us high quality data from the most extreme environments, such as the surface of Venus. Unluckily, they do not operate very well in harsh conditions and that’s why missions have historically only lasted between 23 minutes and 2 hours after a successful landing, before electronics or battery cells got damaged by the planet’s environment.

What we’re up against?

Venus is often referred to as “Earth’s evil twin”, since the two planes share an almost the same size (Venus is 1.1 times Earth, in terms of volume) and gravity (which is around 8.87 m/s2 on the surface). The adjective “evil” comes from the fact that the planet has a high surface temperature (438/482°C), a surface pressure of 93bar and toxic atmosphere (96.5% Carbon Dioxide, 3.5% Nitrogen and 0.015% Sulfur Dioxide). Our task was to find a solution which provides energy to maintain operational a rover or a lander for at least 60 Earth days.

Other interesting data we got from NASA [1]:

Rotation axis inclination: 177.3°

Equatorial radius: 6051.8 km

Surface Gravity: 8.87 m/s^2

Surface pressure: 93 bar

Atmosphere composition:   C0_2 (96.5%), N_2 (3.5%), S0_2

Day duration: 243 Earth days

Year duration: 255 Earth days

Surface temperature: 390°C (med), 482°C (max)

High wind speeds up to 360 km/h 

Thick toxic clouds that scatter almost 90% of the spectrum at high altitudes

The solution

The solution we propose is a system consisting of an orbital part, which stores solar energy through the use of panels, and a lander or rover on the planet's surface, equipped with rechargeable batteries resistant to the extreme environment. The orbiter is in charge of transmitting a directional UV beam towards the lander in order to recharge its batteries and the lander itself through the use of thermoelectric generators picks up the signal and recharges the battery.

For the case study, we assume a demonstration mission with a lander or rover which has an average power consumption of 50W. The system is easily scalable according to power consumption

Energy storage

One of the prominent energy storage technologies for high temperature environments are molten Lithium (Li) and Aluminum (Al) cells (LiAl-FeS2) [2]. They provide a very high energy density and have been extensively developed at Argonne National Laboratory since the early 1990s [3]. Their advantage over other cells technologies is that they operate well between 300°C and 718°C.

For our demonstration lander or rover, it results that a 7.2kg battery (4.67dm3) with a capacity of 928-1000Wh should be enough to last far more over the 94 minutes between one recharge and the following. Considering past missions, the ratio between the weight of the battery and the watts required by the lander is between 0.11kg/W and 0.15kg/W, and in our case we are around the 0.144kg/W. These cells are also easily rechargeable.

The chemistry of the battery includes:

1.     A lithium alloy anode (Li-Al with a melting point of 718°C)

2.     An iron disulfide cathode (FeS2).

3.     A molten salt electrolyte consisting of a eutectic mixture of lithium halides such as LiCl and KCl, which have a low melting point and high ionic conductivity

4.     MgO as the binder, due to its non-reactivity towards lithium at elevated temperatures.

Cathode materials are typically transition metal sulfides, though several other materials (including metal halides) were evaluated, with the critical criterion being their thermal stability.

This cell have has an efficiency of around 50% and a depth of discharge of around 15%. Recent studies have demonstrated the possibility of creating a CR2032 battery with a theoretical capacity of 95%. Lab tests have demonstrated successful operations of the cell in Venus-like surface environment for over 140 days.

Other some important information about the LiAl-FeS2 technology:

•    Operating Temperature range:         300-718°C

•    Open Circuit Voltage (V):                  1.73V

•    Theoretical Specific Energy             490Wh/kg

•    Specific Energy for Cells                    90-130Wh/kg

•    Energy Density for Cells                     150-200Wh/L

•    Cycle Life                                           >1000

Battery integration

Given the estimated volume and weight of the battery pack we opted for a single battery solution. The battery would be housed inside the lander’s body (corrosion and pressure resistant hull). In order to properly insulate and maintain battery efficiency we would need to use some kind of volumetric insulating fiber. Our research shows that the best candidates for insulation are Zirconium Oxide or Ceramic (Alumina monofilament) fabrics, withstanding temperatures of up to 2590°C and 1100°C respectively. In addition, since the mission is mainly a demonstration mission, active thermal control could be envisaged to maintain the optimum battery temperature.

Battery recharging

In order to recharge the lander, we opted for a combined technology exploiting an orbital charging solution. A satellite equipped with solar panels orbits around Venus collecting energy from the sun. The orbit chosen is 350km, to achieve a good compromise between distance and access time (10'). Then, when the satellite is in Line of Sight with the lander, it transmits highly directional and energetic beam towards the lander, in order to recharge batteries. The most promising wavelength to be used is around 240-260nm, i.e. the UVC band. Considering a proper pointing accuracy of the orbiter, the UV beam will reach Venus surface with an efficiency of more than eta=0.9 [4]. The energy is captured using thermoelectric generators mounted on the lander using deployable panels of 3m2 panels. Furthermore, the footprint area of the beam is supposed to be 50 greater than the panels one, in order to compensate potential errors in the pointing. The feasibility and use of UV beams in space has already been demonstrated and utilized [5]. The rays generated are further focused by a system of focalisers. With this solution, 6m2 of solar panels are required to recharge and operate the orbiter itself.

Thermoelectric generator

Thermoelectric generator (TEG), also called a Seebeck generator, is a solid-state device that converts heat flux (temperature differences) directly into electrical energy through a phenomenon called the Seebeck effect (a form of thermoelectric effect). Thermoelectric generators function like heat engines but are less bulky and have no moving parts. However, TEGs are typically more expensive and less efficient. Thermoelectric generators could be used in power plants to convert waste heat into additional electrical power and in automobiles as automotive thermoelectric generators (ATGs) to increase fuel efficiency.

Our approach is using self-extending and temperature resistant thermocouples that will be mounted on the deployable panels. These thermocouples have an efficiency of about 20% [5]. The best performances are observed with nanostructures in bulk thermoelectric that allow for effective phonon scattering of a significant portion of the phonon spectrum, thus greatly improving PbTe (Lead telluride) and energy production.

As a side effect, the panels will also prevent directional UV beam from directly hitting the lander, heating it up. To avoid overheating the thermocouples they can be mounted on some ceramics supports on the top of the lander, eventually adding a thermal heatsink underneath.

Charge controller

Our solution also accounts for a charging circuit. Common Silicon circuitry cannot be used in high temperature environments, so the best substitute are Gallium Arsenide (GaAs) computer boards. GaAs can withstand working temperatures of up to 350°C and support higher frequency clock signals, up to 250 GHz.

NASA Venus Resources https://solarsystem.nasa.gov/news/1519/venus-resources/

Bugga, R. et al. Power Beaming for Long Life Venus Surface Missions, NASA (2019)

Briscoe, J. D. et al. The LiAl/FeS2 battery power source for the future, NASA (1992)

Video sources https://2022.spaceappschallenge.org/challenges/2022-challenges/exploring-venus/details

This experience has been very fulfilling in different aspects. We learned a lot non-only regarding the scientific aspect but also the human one. In fact, our team was composed of people with different courses and passions. Alessandra and Cristiano study aerospace engineering, and Gianmarco and Martina IT engineering. Moreover, we learnt how to trust each other and how to properly cooperate efficiently. We are now ready to take on new adventures with a different self-awareness!

Text references:

[1] NASA Venus Resources https://solarsystem.nasa.gov/news/1519/venus-resources/

[2] Bugga, R. et al. Power Beaming for Long Life Venus Surface Missions, NASA (2019)

[3] Briscoe, J. D. et al. The LiAl/FeS2 battery power source for the future, NASA (1992)

[4] Taylor, F. W. et al.Venus: the atmosphere, climate, surface, interior and near-space environment of an Earth-like planet (2018) Space Science Reviews, 214(1), 1-36

[5] Phipps, C. R. L׳ ADROIT–A spaceborne ultraviolet laser system for space debris clearing (2014), Acta Astronautica, 104(1), 243-255

[6] Biswas, K. et al. High-performance bulk thermoelectrics with all-scale hierarchical architectures, Nature (2012), 489(7416), 414-418

[7] Wertz, J. R. Space mission engineering: the new SMAD (2011), Microcosm Press

Matlab code - sizing electrical system

clc; clear all;

mu_V = 3.24859e14;

R_V = 6051.8e3;

H = 350e3; % Atmosfera venere 250km

R_orb = R_V + H;

V = sqrt(mu_V/R_orb);

T_orb = 2*pi*sqrt(R_orb^3/mu_V) %[s]

T_V = 243*24*60*60; %[s]

alpha = acos(R_V/R_orb);

T_acc=T_orb/(2*pi)*2*alpha

disp( ' Height[1000km] Orbit[min] T_acc[min]')

format bank

[H/1e6 T_orb/60 T_acc/60]

format short

Pland = 50; % P richiesta lander [W]

eff_TC=0.25;

P_received = Pland/eff_TC*T_orb/T_acc; % P durante accesso [W]

eff_atm=0.92; % efficienza UVC (250micrometri) nell'atmosfera

A_ricettore = 3; % Area del ricettore [m2]

Pspec_received = P_received/A_ricettore; %[W/m2]

Aground_trasmessa=A_ricettore*10; % Area su Venere segnale [m2]

Ptrasm=Pspec_received*Aground_trasmessa*eff_atm;

coeff=1.2059; % tengo conto della potenza interna dello spacecraft come percentuale della potenza da trasmettere

rho=asind(R_V/(R_orb));

Te=2*rho*T_orb/360;

Td=T_orb-Te;

Psa=(Ptrasm*coeff*T_acc + (coeff-1)*Ptrasm*(T_orb-T_acc))/T_orb

I=2622; % I venere [W/m2]

Peol = I*cosd(10)*0.35;

A_sa=Psa/Peol

Msa=Psa/300;

%LAnder

eta_batt=0.5;

DOD=0.15;

c=100; %[Wh/kg]

m_b=2.92e-3; %[kg]

C=c*m_b; %[Wh]

N_batt = Pland*(T_orb-T_acc)/3600/eta_batt/DOD/C

Ctot=C*N_batt

C_vol= 150;

Vol=C/C_vol*N_batt %[L]

M_batt=N_batt*m_b

Numero_cicli=60*86400/T_orb;

if Numero_cicli<1000

disp('SI può fare')

else

disp('NON può fare')

end

%((TT/2/pi)^2*mu_V)^(1/3)/1000

%Orbiter

eta_batt=0.95;

DOD=0.25;

c=140; %[Wh/kg]

c_v=

N_cell=1;

C = Ptrasm*T_acc/3600/DOD/eta_batt/N_cell

C_tot=C*N_cell;

M=C_tot/c

V=C_tot/c_v;

%% N=2

Ctot=Ptrasm*(T_acc)/DOD/N/eta_batt

C_vol= 150;

Vol=C/C_vol*N_batt %[L]

M_batt=N_batt*m_b