Orbit Plan and Mission Design for Mars EDL and Surface Exploration Technologies Demonstrator

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1 Trans. JSASS Aerospace Tech. Japan Vol. 4, No. ists3, pp. Pk_9-Pk_5, Orbit Plan and Mission Design for Mars EDL and Surface Exploration Technologies Demonstrator By Naoko OGAWA, ) Misuzu HARUKI, 2) Yoshinori KONDOH, 3) Shuichi MATSUMOTO, 2) Hiroshi TAKEUCHI 4) and Kazuhisa FUJITA 5) ) Space Exploration Innovation Hub Center, Japan Aerospace Exploration Agency, Sagamihara, Japan 2) Research and Development Directorate, Japan Aerospace Exploration Agency, Tsukuba, Japan 3) Human Spaceflight Technology Directorate, Japan Aerospace Exploration Agency, Tsukuba, Japan 4) Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, Sagamihara, Japan 5) Research and Development Directorate, Japan Aerospace Exploration Agency, Chofu, Japan (Received August st, 5) Mars EDL (entry, descent and landing) and surface exploration demonstration working group in Japan Aerospace Exploration Agency (JAXA) has assessed and discussed feasibility of a Martian rover mission to be launched in early s. The primary objectives of this mission are to demonstrate technologies required for EDL and surface exploration of a massive planet with an atmosphere, to investigate Martian geochronology and to search for signs of lives, past or present, and to determine when the ocean was lost in the Martian history. The launch date is targeted in early s in our study. In this paper, we investigate launch opportunities during s and propose several launch windows considering some system requirements. Feasible interplanetary transfer trajectories from Earth to Mars are proposed. Assuming direct entry and following aerodynamic guidance in Martian atmosphere, we connected interplanetary and aerodynamic trajectories so as to land on an aimed point. Precision analysis of orbit determination at the entry and landing is also shown. Key Words: Mars, EDL, Orbit Plan, Orbit Determination, Aerodynamic Guidance. Introduction Mars EDL (entry, descent and landing) and surface exploration demonstration working group in Japan Aerospace Exploration Agency (JAXA) has assessed and discussed feasibility of a Martian rover mission to be launched in early s. ) The primary objectives of this mission are to demonstrate technologies required for EDL (entry, descent and landing) and surface exploration of a massive planet with an atmosphere by driving an autonomous rover, 2) to investigate Martian geochronology and to search for signs of lives, past or present, and to determine when the ocean was lost in the Martian history. The spacecraft system consists of ICM (interplanetary cruise module) and AEM (atmospheric entry module), and AEM is composed of AM (aeroshell module), LM (landing module) and a rover. In this paper, results of our feasibility study on the preliminary trajectory plan, mission design and interface condition for aerodynamic guidance of this mission as of 5 are described. 2. Launch Window Assessment and Trans-Mars Orbit Design In this section, we describe launch window assessment and trans-mars orbit design. 2.. Launch vehicle The nominal squared hyperbolic escape velocity (C 3 ) for transfer from Earth to Mars is around to 2 [km 2 /s 2 ]. In this feasibility study, we assumed H-IIA series capable of such launch requirements as of 5. We also assumed Tanegashima Space Center as the launch site, and coasting flight on the 3- km parking orbit followed by injection into the interplanetary orbit by the upper stage Launch windows Figure shows windows for Earth-Mars transfers between 5 and 25. The blue line shows the sum of hyperbolic excess velocity in Earth departure and Mars arrival. The red line indicates how many times the spacecraft will go around the sun. The launch opportunities which requires rational energy and flight time are,, 22 and. Launch after is reasonable from the viewpoint of development schedule. Among of all, the window in allows us to go to Mars with small energy. Windows in 22 and require more energies because of high declination of launch asymptote and high excess velocity. Thus it can be a good strategy to set as the nominal window, and to regard 22 and as back-up windows. Another strategy for back-up windows can be also feasible by using interplanetary parking orbits followed by Earth gravity assist for trans-mars injection. 3) 2.3. Mission requirements and constraints Next, we assessed each windows, and decided preliminary departure and arrival dates considering the following requirements; Mission Requirements Melas Chasma (29.4 E,.47 S) is the prime candidate of landing sites. Juventae Chasma (29.22 E, 4. S) and Marte Vallis (5.6 E,. N) are also possible. Several possible signs of water have been found around these points, thus we think that they are suitable for life search. Copyright by the Japan Society for Aeronautical and Space Sciences and ISTS. All rights reserved. Pk_9

2 Trans. JSASS Aerospace Tech. Japan Vol. 4, No. ists3 () Vdeparture + Varrival (km/s) Revolution about Sun Launch Date Fig.. Launch windows to Mars between 5 and 25. The blue line shows the sum of hyperbolic excess velocity in Earth departure and Mars arrival. The red line indicates how many times the spacecraft will go around the sun. Time of flight should be within one Earth year, because too long flight may decrease scientific value of this mission. 45-min or more Direct-To-Earth (DTE) communication between Earth (tracking stations in Japan if possible) should be ensured just after landing on Mars, because they have to confirm the success of landing and supervise the lander to acquire power supply and communication as soon as possible. There may be no guarantee for communication relay orbiters dedicated for Japanese Mars missions. The rover should have capability to communicate to Earth for at least 7 sols after landing for sufficient mission activities. System Requirements C 3 values for departure and arrival should be small as much as possible to maximize the probe mass. C 3 on Mars arrival should also be within m 2 /s 2 for thermal protection system on entry. Martian solar longitude (Ls) less than at landing is preferable to avoid mission phase in winter, if the landing site mentioned above is in the southern hemisphere. This is because winter on Mars is too severe thermal condition for the rover. Consecutive 5-days should be ensured for launch, because there is a risk for launch postponement due to weather or other problems. Sun elevation should be about more than 3 degrees just after landing on Mars to ensure sufficient power supply for the rover. Earth distance on Mars arrival should be within. au, required by communication system during cruising. Earth distance 7 sols after landing should be within 2. au for the rover to communicate to Earth by the low gain antenna during the mission phase. Constraint Conditions C 3 on Earth departure should be within 2 m 2 /s 2 constrained by launch vehicle capabilities. Declination of launch asymptote (DLA) should be within ± 6 degrees constrained by the launch site (Tanegashima) and the vehicle Assessment of launch windows DTE communications, Sun elevation, Earth distance and Ls just after landing on Mars are essential requirements in this mission. They depend on Sun, Earth and Mars position relationship, and therefore on departure/arrival dates. We assessed suitable launch and arrival windows to satisfy landing requirements. For example, Fig. 2 shows departure/arrival window candidates superimposed on the porkchop plot for launch opportunities. Red and blue contours are departure and arrival C 3 respectively. Points A to I are candidates Arrival Departure C3 (red) & Arrival C3 (blue) (km2/s2) Sun Elv Departure F C B Ls Earth Dist G A H I E D 2 Departure C3 Arrival C3 A Candidates Earth Ls Dist (au) Fig. 2. Departure/arrival window candidates (A-I) superimposed on the porkchop plot for launch opportunities. Table is a case study result for each windows for landing on Melas Chasma. Note that local time used here is derived by assuming that sol is hours, thus one second is longer than on Earth. You can see that late arrival allows earlier landing in the afternoon with higher Sun and Earth elevation, Pk_

3 N. OGAWA et al.: Orbit Plan and Mission Design for Mars EDL and Surface Exploration Technologies Demonstrator Table. Assessment results for windows. Case Entry date Sun Earth Landing Earthset Sunset Earth Ls Elv [deg] Elv [deg] Local Time Local Time Local Time Dist [au] A :5: :42 7:54 5: B :42: :25 5:49 :. 357 C :49: :3 5:49 : D ::9.5. :25 5:49 : E :7: :59 5:5 7:5. 3 F :6: : 5:5 7:5. 3 G :35: : 5:59 7: H :4: :7 5:54 7:56.4 I :57: : 5:53 7:56.45 which is desirable for DTE communications and power supply just after landing. At the same time, however, late arrival increases Ls and Earth distance, leading to worse communication and season conditions. As the results of trade-off assessment, we selected the candidate I. In an analogous fashion, windows in 22 are also investigated. Sun and Earth elevation and distance showed similar tendency to cases in, because relative angle among Sun, Earth and Mars is almost the same. However, the synodic period of Mars relative to Earth is about days longer than the orbital period of Mars. Therefore, Ls for 22 windows is about 5 degrees larger than that for windows. This means that the arrival will be in winter in the southern hemisphere, which does not meet mission requirements. Thus we presumed Juventae Chasma at low latitudes and Marte Vallis in the northern hemisphere. Conditions for Earth-Mars transfer orbits repeat almost same patterns every 5 years, which is the lowest common multiple of the synodic and orbital periods of Mars. The EDL conditions in 22 is similar to that of Phoenix launched in 7 to land on the northern hemisphere. The summary of assessment is shown on Table 2 as pros and cons of each window. In the case of 22 and launches, more fuel will be required because we have to launch the vehicle toward southeast. Moreover, 22 and are severe for exploration of the southern hemisphere Orbit design results Considering assessment results mentioned above, we designed trans-mars orbits. For example, here we describe some results for launch as follows. Note that these are preliminary results presuming ballistic flights, and no multibody dynamics, perturbations, trajectory correction maneuvers are regarded so far. Launch date from Tanegashima Space Center will be around :59 a.m. JST, 4th Aug.. The vehicle will be launched toward east with the launch azimuth of 9 degrees, and after 64- min coasting on the parking orbit of 3-km altitude, injected into the departure orbit by burn of the second stage followed by the probe separation. The long coast is adopted to prevent from umbra after separation. The ground track during the launch is shown in Fig. 3. Figures 4-7 shows the profiles of Mars transfer orbit and visible passes from Usuda Deep Space Center, respectively. Just after the launch the spacecraft is visible in midnight, but the pass will become earlier gradually. Latitude [deg] Y [km] (J. Ecliptic) Separation: --4 7:4:7 UTC Launch: --4 5:59:57 UTC Longitude [deg] Fig. 3. Ground track examples for the launch in. 3 x 2 Arrival Sun S/C Earth Mars Departure X [km] (J. Ecliptic) x Fig. 4. Mars transfer orbit in (Black: Spacecraft, Blue: Earth, Red: Mars). 3. Interface Condition for Trajectory inside Martian Atmosphere We assume the demonstrator to perform direct entry from the interplanetary orbit to the Martian atmosphere. Aerodynamic guidance should be executed inside the atmosphere, while ballistic flight is assumed in the interplanetary space. Trajectory in each phase is computed by using independent algorithm, but two trajectories should be connected smoothly. Thus we de- Pk_

4 Trans. JSASS Aerospace Tech. Japan Vol. 4, No. ists3 () Table 2. Launch window assessment summary. Window 22 Departure Arrival TOF Ls on Arrival Departure V (km/s) Arrival V (km/s) DLA (deg) RLA (deg) Launch Capability Good Fair Fair Injection Mass (t).4.. Earth Distance on Arrival (au) Sun Distance on Arrival (au) Misc Nominal Distance (AU) Angle (deg) Arrival (2-2-27) Cruise --2 Fig. 5. Cruise Mission Distance from Earth Ls Distance from Sun Date Sun and Earth distance in Launch. Arrival (2-2-27) 2-4- SPE Mission Date SEP Fig. 6. Sun-Earth-Probe angle and Sun-Probe-Earth angle in Launch Ls (deg) Visible from Japan UTC Fig. 7. Visibility from Usuda Deep Space Center in Launch. fined the entry point as the interface between two trajectories in the interplanetary space and in the air, based on the preliminary study results for aerodynamic guidance; Altitude: 5 km Flight Path Angle: 7 degrees As the feasibility study, we assumed the landing target Melas Chasma at longitude 29.4 degrees east and latitude.47 degrees south, and the parachute deployment point to be -km above the landing target. The following results are derived from a case study assuming the launch in. First, given the ephemeris of the departure and arrival, we solve the Lambert problem to obtain the magnitude V, right ascension α and the declination δ of the approaching hyperbolic excess velocity. We can adjust the injection direction by changing the phase angle θ on the B-plane for the approach to Mars. The altitude h and the flight path angle ϕ are given as mentioned above. With respect to the arbitrary θ, we can acquire a trajectory to meet given V, α, δ, h and ϕ by iterative calculation, and thus the longitude, the inertial velocity and the flight azimuth of the interface point on this trajectory. Therefore, we created a look-up table showing the longitude, the inertial velocity and the flight azimuth of the interface point as we changed θ from 9 to 9 degrees (we assumed entry along the Mars rotation). Figures - show the relation among Pk_

5 N. OGAWA et al.: Orbit Plan and Mission Design for Mars EDL and Surface Exploration Technologies Demonstrator the longitude, the inertial velocity and the flight azimuth of the interface point in this look-up table. Inertial Velocity at I/F Point (km/s) Flight Azimuth at I/F Point (deg) Fig.. Inertial velocity with respect to the flight azimuth on the interface point in the launch case. Flight Azimuth at I/F Point (deg) Geographic Latitude of I/F Point (deg) Fig. 9. Flight azimuth with respect to the geographic latitude on the interface point in the launch case. Table 3. Position and velocity at the entry interface point. Parameter Unit Value Date (TDB) :57: Altitude km 5. Geographic Latitude degn.649 Longitude dege Inertial Velocity m/sec 55.4 Flight Path Angle of Inertial deg 7. Velocity Flight Azimuth of Inertial Velocity deg X (Mars-centered ICRF) km 9.7 Y (Mars-centered ICRF) km Z (Mars-centered ICRF) km VX (Mars-centered ICRF) km/s 3.57 VY (Mars-centered ICRF) km/s 4.75 VZ (Mars-centered ICRF) km/s.527 Z (km) x B-Plane Ax Equinox S Inertial Velocity at I/F Point (km/s) Geographic Latitude of I/F Point (deg) Fig.. Inertial velocity with respect to geographic latitude on the interface point in the launch case. From this table, we can obtain the longitude and latitude of the interface point suitable for landing of the desired target by interpolation. We adjusted the connection between the trajectories again to refine the position and velocity on the interface point shown in Table 3. Figure shows the Mars approach trajectory. The spacecraft will enter the Martian atmosphere at :57 UTC 27th Feb. 2 at the altitude of 5 km and the south latitude of.6 degrees. The entry velocity is around 5.5 km/s. The local time on the landing site (Melas Chasma) will be around 2 o clock p.m. The landing site is visible from the tracking stations in Japan, and vice versa. -.5 x 4 Fig.. Earth Y (km) - Sun T R x 4 X (km) Mars approach trajectory seen from Martian north pole. 4. Orbit Determination and Aerodynamic Guidance Before EDL, precise orbit determination using Delta-DOR will be performed. An estimation result is shown in Fig.. It is indicated that the orbit can be determined with about -km precision. Using orbit determination results, a Monte Carlo simulation for aeroassisted guidance flight at EDL was performed using real-time prediction guidance 4) and Mars-GRAM (Mars Global Reference Atmospheric Model). Parameters of the entry capsule was assumed as shown in Table 4. Using the atmospheric model and the nominal trajectory analysis tool (POST), 5) some iteration process between interplanetary trajectory design and aeroassisted flight path determined entry interface point condition as shown in Table 5. As for error sources of the Monte Carlo simulation, we assumed errors of position and velocity as initial state errors, errors of position, velocity and attitude Pk_3

6 Trans. JSASS Aerospace Tech. Japan Vol. 4, No. ists3 () Longitude(deg) (km) Latitude (deg) (km) Fig.. An orbit determination result at the interface point in the case. Table 4. Parameters of the entry capsule used in the Monte Carlo simulation of aeroassisted guidance. Parameter Unit Value Mass kg 629. Ballistic coefficient 2.3 Drag coefficient.44 Lift coefficient.29 L/D. Trim angle deg 3.9 Diameter m 2.6 Area m Nose radius m.65 Table 5. Parameters of the atmospheric entry used in the Monte Carlo simulation of aeroassisted guidance. Parameter Unit Value Altitude km Geographic latitude degn Longitude dege Inertial velocity m/s Flight path angle deg -7.4 Flight azimuth deg as initial navigation errors, atmospheric density error and orbit determination errors. The simulation result in Fig. 3 implies that the capsule can be guided with sufficient accuracy to the parachute deployment point. The size of the error ellipse is about km 6 km. According to the wind analysis, east-west and north-south error by the wind are assumed to be km and 3 km, respectively. Thus the final error ellipse for landing will be about km 4 km, which is rational and comparative with the assumed error ellipse for Mars ( km 4 km). 6) Because the rover has high mobility on Mars, the accuracy of landing is sufficient for exploration of the landing site candidates. 5. Summary We described a preliminary trajectory plan for the Mars EDL demonstrator mission. Feasible interplanetary transfer trajectories from Earth to Mars were proposed considering mission requirements. Precision analysis of orbit determination before the entry and guided flight in Martian atmosphere were also shown. In future works, the operation analysis, consideration Geographic Latitude (deg) Parachute Deployment Points Variance Ellipse (3-sigma) Landing Target Radius of 5 km Longitude (deg) Fig. 3. A Monte Carlo simulation result for guided reentry flight in Martian atmosphere in the case. of planetary protection, contingency and backup plans are to be discussed. References ) Fujita, K., Ishigami, G., Hatakenaka, R., Takai, M., Toyota, H., Ogawa, N., Haruki, M., Takeuchi, H., Nonomura, T., Yamada, K., Takayanagi, H., Ozawa, T., Matsuyama, S., Oyama, A., Yamagishi, A., Kameda, S., Miyamoto, H. and Satoh, T.: Japan s Mars Rover Mission - System Design & Development Status, Proceedings of The 3th International Symposium on Space Technology and Science, 5-k-37, 5. 2) Ishigami, G., Fujita, K., Hatakenaka, R., Toyota, H., Sato, T., Takai, M. and Nonomura, T.: Mission Scope Definition and Preliminarily Design Study of Mars Surface Exploration Rover, Proceedings of The 3th International Symposium on Space Technology and Science, 5-k-39, 5. 3) Ogawa, N., Mimasu, Y., Tanaka, K., Yamaguchi, T., Fujita, K., Narita, S. and Kawaguchi, J.: Earth Revolution Synchronous Orbits and Aero-Gravity Assists to Enhance Capabilities for Interplanetary Missions by Sub-Payload Spacecraft, Advances in the Astronautical Sciences, 46(3), pp ) Matsumoto, S., Kondoh, Y., Suzuki, Y., Yamamoto, H., Kobayashi, S. and Motoyama, N.: Accurate Real-Time Prediction Guidance Using Numerical Integration for Reentry Spacecraft, AIAA Paper , 3. 5) Brauer, G. L., Cornick, D. E. and Stevenson, R.: Capabilities and Ap- Pk_4

7 N. OGAWA et al.: Orbit Plan and Mission Design for Mars EDL and Surface Exploration Technologies Demonstrator plications of the Program to Optimize Simulated Trajectories (POST), Technical Report, NASA CR-277, ) Chen, A., Hines, E., Otero, R., Stehura, A. and Villar, G.: Mars Entry, Descent, and Landing System Overview, th International Planetary Probe Workshop, 4. Pk_5