LEMFEV
LONG ENDURANCE MARS EXPLORATION FLYING VEHICLE
Designing a Martian UAV
Here, we share our engineering and scientific findings and reflect on our design-related choices
What atmospheric conditions are expected for the LEMFEV on Mars?
The Martian atmosphere is composed mainly of carbon dioxide (96% of CO2). The average surface pressure of the Martian atmosphere is about 1% of Earth's sea level pressure. The surface temperatures on Mars vary from -140 °C to +20 °C. Furthermore, strong winds (up to 5 m/s during the daytime) and turbulence can degrade aircraft efficiency and stability. The density at the surface of Mars is as low as that at an altitude of approximately 32 km above the Earth's surface. Figure 1 shows the distribution of nominal atmospheric density with altitude for Mars as compared to the data for the Earth's standard atmosphere.
Figure 1 - Mars Nominal Atmosphere [1] and Earth Standard Atmosphere [2]: Density vs Altitude
Density on Mars is highly variable depending on seasonal and daily cycles. During the daytime, it can fluctuate between 0.014 and 0.020 kg/m3.
The speed of sound on Mars is also lower than that on Earth due to the low temperature and different composition of the atmosphere; therefore, compressibility effects are more easily triggered. Figure 2 shows the dependence of the speed of sound on the altitude for Mars and Earth.
Figure 2 - Mars Nominal Atmosphere [1] and Earth Standard Atmosphere [2]:
Speed of Sound vs Altitude
Due to these specific environmental conditions, as well as unknown terrain elevation, limited knowledge of wind speed, dust storms, and rugged terrain, a Mars flying vehicle should possess some features different from those of an Earth aircraft.
References:
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NASA Technical Reports Server (NTRS) 20010056680: Mars Global Reference Atmospheric Model 2001 Version (Mars-GRAM 2001): Users Guide.
Propulsion system capability is the key element in establishing the aircraft feasibility and reaching the desired flight envelope. This is especially true for an aircraft that is to fly on other planets, including Mars. The specific Mars environment, as well as the launch from Earth and transit in deep space, produce significant obstacles to airborne platform performance.
Without the ability to refuel, mission duration is to be limited by the amount of energy that can be carried onboard the aircraft. Therefore, the more efficient and lightweight the propulsion system is, the longer the mission.
Low atmospheric density assumes issues with generating thrust. The lower the atmospheric density, the less mass available for momentum transfer and the less thrust that can be generated for a given propulsion system. Therefore, a propeller designed for Mars conditions would be larger than a typical Earth propeller to generate the same amount of thrust.
A lack of appreciable amounts of atmospheric oxygen presents another propulsion issue for a Martian aircraft. Most Earth powered aircraft use air-breathing propulsion systems, relying on the oxygen from incoming air as an oxidizer to release energy stored in the on-board fuel. This approach is not viable in the mainly carbon dioxide Martian atmosphere. Numerous concepts exist for non-airbreathing propulsion systems. The common drawback of the majority of these systems is that they have much higher fuel consumption and additional mass and complexity compared to conventional airplane propulsion systems.
Propulsion options potentially feasible for Martian conditions include:
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no propulsion (glider),
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rocket (liquid or solid), and
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propeller (driven by various power sources).
A glider is attractive because for a Martian glider, the mass, cost, and risk associated with the propulsion system are eliminated. The obvious downside to a glider is related to the fact that its range is determined by the glide slope and the starting altitude. On the one side, the higher the starting altitude, the greater the range. On the other side, the operating altitude of the on-board scientific instruments may be limited.
Rocket propulsion is believed to be one of the lowest risk options for a powered Mars aircraft. However, the need for controlled thrust over a comparatively long time period greatly increases the complexity of a solid rocket system.
In the Earth conditions, a propeller is a more efficient means of generating thrust than a rocket for a low-speed, high-altitude airplane. For Mars application, a number of propeller-based propulsion systems were considered as well.
Options considered for driving the propeller in Martian conditions included:
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an electric motor powered by batteries (consisting of batteries, propeller, gearbox, electric motor),
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an electric motor powered by a fuel cell, and
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the Akkerman type hydrazine engine (consisting of hydrazine engine, piston expander).
On Earth, the search for clean power systems has led to the development of alternatives to the traditional lead–acid battery. One of them is a fuel cell, which is a chemical and mechanical device to convert chemical energy stored in a source fuel into electrical energy without the need to burn the fuel. This is potentially highly efficient, has almost harmless emissions, and is quiet (which is of primary importance for Earth applications).
Low Density, Low Reynolds Number, High Mach Number - Aerodynamic Design Options for a Martian UAV
In a thin atmosphere, wing loading has to be small. On the other side, on Mars, the airplane weight is only approximately 38% of what it would be on Earth due to the lower gravity. Nevertheless, for a Mars aircraft, the airframe and system masses must be as low as possible in order to achieve the required low wing loading at the least possible wing area.
For the given aircraft weight, the low atmospheric density is associated with a high cruise Mach number (greater than 0.5), but at low flight Reynolds number (of the order of 50 000, for the previous projects). Since the speed of sound on Mars is lower than on Earth, the aircraft will suffer transonic aerodynamic effects at a comparatively low speed. Therefore, the measures should be taken to ensure the kinetic energy level of the boundary layer high enough to resist premature separation; on the other side, the flow acceleration on the surface of the airframe should be limited to avoid shock waves.
Aerodynamic layout options for Mars aircraft proposed since 1970-s include:
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conventional wing-tail layout with a high aspect-ratio straight wing and winglets (Colozza, 1990), (Smith, et al., 2000), (Gasbarre, 2003), (Walker, 2008);
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flying wing with wingtip vertical stabilizers and with the vertical stabilizer on the fuselage.
Flying wing layout is promising because of its potential of eliminating the tail parasite drag, thereby obtaining a gain in the aerodynamic efficiency. Also, it offers propulsion integration and folding simplicity (one fold for each wing). However, the trim drag of a flying wing should be relatively high. Probably, for this reason, most of Mars aircraft feature a conventional wing-tail layout.
Below, some historical data on the fixed-wing Mars aircraft projects is presented.
Table 1 - Historical Data on the Existing Fixed-Wing Mars Aircraft Projects
Although Table 1 doesn’t include the data on the aircraft with too few known parameters or extreme parameter values (like (Colozza, 1990), (Aguirre, Casado, Chamie, & Zha, 2007)), the wing span, area, as well as the take-off and payload masses change in a wide range. This might be attributed to the variation in scientific mission, technical advances, as well as to the uncertainty in the input design data. The average value of non-dimensional design parameters, excluding too extreme values, are presented in Table 2.
Table 2 - Averaged historical values of some Martian UAV parameters
An alternative view on possible means of achieving flight in the Martian atmosphere includes flapping wings concept (Peeters, et al., 2008), and even a Mars bee (Bluman, Kang, Landrum, Fahimi, & Mesmer, 2017).
It has been suggested that flapping aerial vehicles may provide an attractive solution for the exploration of Mars, due to low gravity and low Reynolds numbers on Mars.
In (Peeters, et al., 2008), it was reported that the demonstrator with a total weight of 17 gr and capable of flying for 12 min with onboard energy storage and a pinhole camera payload was tested. In Mars conditions, the equivalent vehicle would have the mass of 20 gr and range of 10-15 km with onboard solar cell recharging of the energy storage subsystem and a similar scientific payload.
For the LEMFEV, we selected two wing layouts: the conventional wing-tail and the boxwing layout. The analysis will show which one matches better the Martian scientific mission.
References Aguirre, J., Casado, V., Chamie, N., & Zha, G. (2007). Mars Intelligent Reconnaissance Aerial and Ground Explorer (MIRAGE). Bluman, J., Kang, C., Landrum, D., Fahimi, F., & Mesmer, B. (2017). Marsbee - Can a Bee Fly on Mars? Colozza, A. (1990). Preliminary Design of a Long-Endurance Mars Aircraft. Gasbarre, J. R. (2003). Preliminary design and analysis of the ARES atmospheric flight vehicle thermal control system. SAE Technical Paper. Peeters, B., Mulder, J., Kraft, S., Leijtens, J., Zegers, T., Lentink, D., & Lan, N. (2008). EXOFLY: A FLAPPING WINGED AEROBOT FOR AUTONOMOUS FLIGHT IN MARS ATMOSPHERE. Smith, S., Hahn, A., Johnson, W., Kinney, D., Pollitt, J., & Reuther, J. (2000). The Design of the Canyon Flyer, An Airplane for Mars Exploration. Walker, D. (2008). Preliminary Design, Flight Simulation, and Task Evaluation of a Mars Airplane.
MATLAB code for the design and analysis of the LEMFEV
Fig. 1 shows the structure of the MATLAB software developed for sizing and analysis of the WT1 (wing-tail electric) and WT3 (wing-tail rocket-based) LEMFEV configurations. The main code is based on constraint analysis and the unity equation. It takes data from the input block (solar irradiance, atmospheric parameters, baseline airfoil aerodynamic properties, design flight conditions, engine, systems, and equipment specifications). The code produces the aircraft and its component masses, as well as the wing area, required thrust, flight performance, rational airfoil and its aerodynamic properties, tail geometry, and a limited stability analysis for the relevant operating conditions. In this context, ‘rational airfoil’ means the airfoil that features the greatest value of a measure of quality (discussed below) at the estimated design conditions.
Figure 1 - The MATLAB code structure developed for the LEMFEV project
In Fig. 1, W = weight [N]; E = endurance [min]; d = diameter [m]; S = wing area [m2]; b = wing span [m]; AR= b^2/S = wing aspect ratio [-]; T = thrust [N]; V_cr = cruise speed [m/s]; i = iteration number; P = power [W]; V_(HT,VT) = horizontal and vertical tail volume coefficient [-]; X_(AC,CG) = x-coordinate of aerodynamic center and center of gravity reduced by the length of mean aerodynamic chord (MAC) [-]; R = the ratio of solar cell area to wing area; m = mass [kg].
The Constraint Diagram
The constraint diagram shown in Fig. 2 was used to establish the optimum wing loading W/S and thrust loading T/W for the WT1 and WT3 configurations.
Fig. 2 WT1 and WT3 Constraint Diagram: AR 3, Payload 7 kg.
In Fig. 2, V_st= stalling velocity; C_(L,max) = maximum lift coefficient.
The considered design cases included:
T/W for a service ceiling (rate of climb at ceiling 0.508 m/s);
T/W for a desired rate of climb;
T/W for a desired cruise airspeed;
T/W for a level constant-velocity turn.
Also, the diagram shows what maximum lift coefficient the UAV’s airfoil must have to satisfy the given stalling velocity and wing loading requirements. Higher wing loadings and lower stalling velocities call for exceedingly high maximum lift coefficients.
For reference, the diagram shows the design stalling velocity limit (which is the function of the UAV’s weight and maximum lift coefficient), as well as the wing loading corresponding to the available load factor (maximum lift to weight ratio) for the given Mach number limitation (0.8).
The optimal design point permits establishing the wing and thrust loadings at which all the requirements are met. At the design stall limit of 52 m/s, the maximum allowable wing loading is approximately 25 N/m^2. The design thrust loading is 0.24. The iterative solution of the unity equation yields the optimized value of the airplane’s gross mass, hence the design wing area and thrust.
One of the key aspects of a solar UAV design is the selection of the battery and solar cells.
At present, a new type of sulfur cathode (Li-S) battery is being introduced into technical systems with high energy storage capacity demand. This technology begins to replace lithium-ion batteries.
The main advantages of Li-S over Li-ion are the following:
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higher values of specific energy storage capacity, which for test samples may reach up to 500 Wh/kg ;
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protection against overcharging is not required;
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wider operating temperature range;
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safety in case of mechanical damage.
The main disadvantage is a large degradation due to charge-discharge cycles compared to Li-ion batteries. Also, this technology is still expensive for the commercial replacement of Li-ion batteries.
The Energy Balance Profile
The energy balance profile compares the total energy required for flight, navigation, and scientific instruments to the energy available from the battery over a 24-hour cycle. It acсounts for the need to reduce the voltage for avionics and payload, the charge and discharge efficiency of the battery used for the overnight flight, and the efficiency of solar cells and the maximum power point tracker.
Fig. 4 shows the energy balance profile for the WT1 configuration. The excess electrical power refers to the power available for charging the battery. Under estimated conditions, the UAV’s battery can be charged during the day, so an overnight flight is possible.
Fig. 4 WT1 configuration. Cruise energy balance profile: AR 3, altitude 1000 m, Re 1.18e+05, Cl 0.85
WT1 and WT3 Specifications
In Table 1, the WT1 and WT3 specifications are given. These include the flight and design conditions, weight specification, aircraft geometric and performance characteristics for a payload mass of 7 kg and a wing aspect ratio of 3.
Table 1 - WT1 vs WT3 Specifications
The geometries of the WT1 and WT3 configurations are compared in Fig. 5.
Figure 5 - WT1 and WT3 configurations, top and side views
The optimized versions of the two configurations feature the same wing and tail airfoils, the same wing and tail planforms, and different dimensions and weights. The two aircraft have similar cruise speeds and are designed to carry the same payload. The need to place the rocket engine, fuel tank, and pressurized vessel in the fuselage renders the WT3 configuration more challenging to size. The endurance of the rocket-based aircraft is limited by the fuel tank volume and fuel mass; however, the operating area of this aircraft is not restricted to the regions, seasons, and days with high solar irradiance. Therefore, for example, it can explore the depths of crators and canyons and, if a single soft landing option is possible, it can also serve as a stationary platform. Theoretically, the endurance of the solar UAV is unlimited; the drawback is that it can only operate under specific atmospheric conditions. Since the solar irradiance depends on the geographic location, aerocentric longitude, and albedo, the success of the mission performed by a solar UAV is highly uncertain. A suitable scientific mission for a solar single-flight UAV may be the measurement of atmospheric parameters to obtain turbulent and radiative fluxes over the lowest 2–10 km of the atmosphere. This will allow us to to expand the geographical and temporal coverage of measurements currently available for the planetary boundary layer of Mars.
The theoretical day-night flight of a solar Martian airplane became possible due to advancements in battery technology. With the battery specific energy being lower than 400 Wh/kg, at Martian solar irradiation, the battery cannot be charged during the daytime to ensure an overnight flight.
The Earth-Mars flight will follow the Homan half-ellipse, and the flight will last for 259 days. On the flight path, streams of charged particles of galactic cosmic rays (GCL) and solar cosmic rays (SCL) will impact the equipment and devices of the flight module with the payload (Martian UAV). We analyzed the expected radiation loading using the dynamic GCL model and the probabilistic SCL model, modified to take into account the radial dependence of SCL fluxes as the spacecraft moves away from the Sun. Since the SCL model is probabilistic in nature, when calculating the fluxes of solar cosmic ray particles and their contribution to the absorbed dose, a probability of 0.1 was set (this means that exceeding the given values of fluxes and absorbed doses during flight is possible only in 10% of cases).
The amount of the local absorbed dose depends on the location of the device in the spacecraft, its shielding by other devices and structural elements. Radioisotope heat sources (RHS) can be used on board the flight module.
The amount of absorbed doses in the equipment and devices on board depends on the distance to the RHS.
After landing, the Martian UAV is expected to operate for more than 3 Earth years – the calculations assumed an active service life of 1,100 days. The thickness of the Martian atmosphere is estimated at 20 g/cm^2. Absorbed doses during the operation of the UAV in the atmosphere and on the surface of Mars will amount to about 14 rad (Si).
In analysis, we have adopted the following indicators of the resistance of microcircuits to random single effects: the saturation cross section of the effect and the magnitude of the threshold linear energy transfer (LET) of the effect.
Thus, for the Martian aircraft, these parameters are as follows:
For catastrophic failures:
- saturation cross section: 10^2 cm^2/chip;
- threshold LET: 20 MeV*cm^2/mg.
For reversible failures:
- saturation cross section: 10^2 cm^2/chip;
- threshold LET: 1.5 MeV*cm^2/mg.
These values of threshold LPE and saturation cross sections are inherent in electrical and radio products (ERP) with low fault tolerance.The probability of failure-free operation of such an ERP in 1100 days will be 0.992. The number of failures during an extreme solar event will be ~ 1 failure in two days.