Mitigation of Restrictions in Planetary Missions by using Interplanetary Parking Orbits and Aeroassist

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1 Mitigation of Restrictions in Planetary Missions by using Interplanetary Parking Orbits and Aeroassist Naoko Ogawa, Yuya Mimasu, Kazuhisa Fujita, Hiroshi Takeuchi 3, Keita Tanaka 4, Shinichiro Narita and Jun ichiro Kawaguchi 3 JAXA Space Exploration Center, Japan Aerospace Exploration Agency Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency 3 Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency 4 Graduate School of Engineering, University of Tokyo Abstract: In general, planetary missions have many restrictions such as few launch opportunities and mass limit. We propose a way to enhance opportunities of planetary missions by trajectory design incorporating interplanetary parking orbits and aeroassist. A case study of our proposal is described : Introduction Progress of capability in launch vehicles has offered us more and more opportunities for sub-payloads or piggy-back satellites. Now it is common that one launch vehicle carries several satellites other than the main payload. Even chances for interplanetary missions have been opened to subpayloads in the case of planetary explorers as the main payloads. For example, the world-first solar power sail demonstrator IKAROS [] and the world-first university-made deep space satellite UNITEC- [] were launched as sub-payloads of Venus Climate Orbiter Akatsuki [3] in. Generally sub-payload spacecraft are restricted from free mission design, however, because of the higher priority of the main mission. Weight, schedule, separation timing and final orbit of the sub-payloads deeply depend on those of the main spacecraft. Especially in interplanetary missions, the destination must be almost fixed. This restriction would be an severe obstacle for sub-payloads aiming independent and free missions, though it would be a rare opportunity of a deep space flight for the sub-payloads. One promising scheme to expand mission opportunities for sub-payloads is to introduce Earth revolution synchronous orbits into the mission sequence. These orbits have periodicity synchronous with revolution of Earth, and the spacecraft will encounter with Earth periodically regardless of its hyperbolic velocity with respect to Earth [4, 5]. By throwing main and sub payloads toward different directions at the gravity assist in the encounter with Earth, we can guide the sub-payloads to trajectories completely independent from the main payload, as shown schematically in Fig.. It also enables us to increase chances for transfer windows, because the sub-payloads can choose the best timing to transfer by waiting on synchronous orbits. () Direct Launch Main Injection into Interplanetary Orbit Sub Earth Revolution Synchronous Orbit () Interplanetary Parking along Earth Revolution Synchronous Orbit Re-Encounter with Earth after.5, or.4 years (3) Injection into Different Orbits by Earth Gravity Assist Different α inf δ inf (V inf ) Gravity Assist Option: Aero-Gravity Assist Fig. : A concept of the proposed scheme to enhance interplanetary missions for a sub-payload by Earth revolution synchronous orbits. Controllable variables in the proposed scheme with Earth revolution synchronous orbits are the right ascension and the declination of the Earth departure velocity. Additionally, the departure velocity itself will also be controlled if we introduce Earth aero-gravity assists. It will greatly enhance interplanetary mission opportunities for sub-payloads as shown in Fig..

2 8 6 Decrease V inf Dec [deg] 4 - Aero-Gravity Assist Fig. : A concept of the proposed scheme to enhance interplanetary missions for a sub-payload by Earth aero-gravity assists. In this paper we propose a strategy for sub-payload spacecraft to extend possibility of interplanetary exploration without interrupting the mission by the main spacecraft. We show a case study for a possible mission design of a sub-payload bound for Mars and a preliminary mission sequence. Earth Revolution Synchronous Orbits Earth revolution synchronous orbit is an orbit which has periodicity synchronous with revolution of Earth. There are three types of Earth revolution synchronous orbits: (A) oneyear loop orbit, (B).5-year transfer orbit, (C) singular orbits with irrational periods of about.4-year,.4-year... loops. Figure 3 illustrates various types of Earth revolution synchronous orbits. One-year orbits have the best freedom of design, because requirement for the orbits is just to have the semi-major axis of AU. Inclination of the heliocentric orbit, or departure declination after the gravity assist can be chosen freely. Figure 4 shows an example for pairs of right ascension and declination of the outgoing hyperbolic excess velocities for one-year loop Earth revolution synchronous orbits in December. We can select preferable declination on the solid line according to mission requirements. Moreover, the energy sensitivity tends to be low, and only a small delta-v is needed for trajectory correction maneuvers..5-year orbits additionally require the same mean anomaly at the gravity assist, and the resulting orbits are obtained by changing inclination of the Earth orbit. Singular orbits are derived from in-planar Hill s equation and its solution of the period is.4,.4,... and so on. In all cases, orbits has two options: to go inbound first after the gravity assist, or outbound first. This scheme can easily be applied to other planets. Such variation of synchronous orbits are useful to ensure timing buffers for transfers to destinations as a kind of heliocentric phasing orbits, or to increase additive launch windows. It is required for the main payload to adopt Earth revolution synchronous orbits and gravity assists before its cruise to the RA [deg] Fig. 4: An example for pairs of right ascension and declination of the outgoing hyperbolic excess velocities for one-year loop Earth revolution synchronous orbits. destination, which can be an advantage also for the main payload to enhance mission capabilities by additive windows or schedule margins. 3 Aero-Gravity Assists Another advantage of using Earth revolution synchronous orbits is possibility to decrease departure C3 energy by means of aerobraking at synchronized swingbys, or aero-gravity assist [6,7], which can lead to the reduction of insertion delta-v at the destination planet. In dual-launch missions, the departure C3 is designed to be suitable for the main payload, which may be sometimes too large for sub-payloads targeting more accessible destinations. Generally sub-payloads are subject to severe restrictions in their masses, thus forced to carry only small amount of propellants. It is difficult for sub-payloads to generate large delta-v at orbit insertion or mid-course maneuvers. One of effective scheme to reduce the orbit velocity is aero-gravity assist. Aero-gravity assist at Earth is a kind of aerobraking, or it can be regarded as a de-powered swingby. By passing through Earth atmosphere, the spacecraft can be decelerated. Aerobraking technique was first achieved by Hiten (MUSES- A) on a cis-lunar orbit [8], and then several spacecraft have utilized it to lower the apoapsis over Venus or Mars. At the earth gravity assist during synchronous orbits, hyperbolic flight through the upper atmosphere at about 9 to -km altitude will generate aerodynamic deceleration. Precise orbit determination like delta differential one-way range (delta- DOR) will be effective for increasing maneuver accuracy. One major difference between aero-gravity assist and usual swingby without aerobraking is that the extent of deceleration and the perigee altitude are coupled and dependent to each other in aero-gravity assists. In order to obtain a certain deceleration under thermal conditions, we have to fix the perigee altitude and the resulting outgoing direction. Thus, we solve

3 Z.5rev Northward.5rev Southward rev Outbound rev Inbound.4rev Outbound Fig. 3: Various types of Earth revolution synchronous orbits..4rev Inbound this problem by allocating a one-year Earth revolution synchronous orbit with degrees of freedom in orbits just after the aero-gravity assist so that we can bring back the spacecraft to Earth regardless of the outgoing direction. For simplicity, in the design process, we assume the deceleration as an impulse delta-v at the perigee, and obtain right ascension and declination of the outgoing V inf so as to constitute a one-year synchronous orbit and to pass the rational altitude at the perigee for the given incoming V inf. Low sensitivity of one-year synchronous orbits can also absorb orbit errors or perturbations after passing through the atmosphere efficiently. An example of aero-gravity assist is illustrated on Fig. 5. Note that the extent of deceleration in this example does not consider actual tolerance of existing thermal protection systems, and it might be too deep. Deceleration in the actual mission will be decided via tradeoff of thermal impact at Earth and the necessary deceleration at the target planet x 4 Arrival 4.473km/s R x 4 Departure Y x 4 X.87km/s B T B Perigee 8km Perigee Velocity.98km/s->.456km/s ( ΔV = -55m/s) Fig. 5: An example of aero-gravity assist by Earth. 4 A Case Study Here we show a preliminary case study of the proposed strategy. Let us assume that our spacecraft is a sub-payload of Hayabusa-, and that our destination is Mars or Venus. Hayabusa- is an asteroid sample-return spacecraft whose destination candidate is Asteroid 999 JU3 [9], and now under development. There are several candidates of launch windows in 4. The main spacecraft will utilize the Electric Delta-V S Earth Gravity Assist (EDVEGA) scheme after the launch []. EDVEGA, or Solar Electric Propulsion Leverage is a technique to garner the orbital energy by combination of electric propulsion and Earth revolution synchronous orbits []. By applying this scheme, Hayabusa- will follow a.4,. or.5-year loop orbit with the low-thrust propulsion just after its launch till the Earth swingby in the winter of 5. It means that the sub-payload will also be injected into the synchronous orbit together with the main spacecraft. Acceleration of the main spacecraft is so small that the trajectories of the two spacecraft are very close and the sub-payload will also encounter Earth in the winter of 5. The gravitational slingshot can throw two spacecraft toward different direction and the main spacecraft will be headed to the asteroid. 4. Trajectory Design Examples for Mars First, we investigate a case for Mars. After the Earth gravity assist in the winter of 5, when and to what direction to transfer to Mars should be decided. Figure 6 shows departure windows from Earth to Mars during 6-9. It indicates that there are two preferable windows, 6 and 8. We investigated whether launch windows can connect its orbit to Mars transfer via Earth gravity assists. It was revealed that some windows in 4 cannot be connected to the leg for the Mars transfer orbit. Thus we decided that the sub-payload should wait for a next chance of transfer to Mars while tracing synchronous trajectories again. For the next departure window will be the spring in 8, we have to wait for two and a half years. Combination of several Earth revolution synchronous trajectories and gravity assists can be applied. For example, we can combine one-year and.4-year inbound Earth revolution synchronous orbits as illustrated in Fig We confirmed that all the launch windows can be connected to Mars transfer orbit in 8 by applying this technique. Aero-gravity assists can be also performed before or after one-year loops, which will reduce delta-v at Mars orbit insertion. Taking the points discussed above into account, we designed a preliminary trajectory for each launch window. Fig- 3

4 9 8 Total Vinf Window Window 7 6 Arrival Vinf 5 4 Vtot Vdep Varr 3 Departure Vinf Revolution N Fig. 6: Departure windows from Earth to Mars..5 x Swingby- 6. Swingby- 8.5 Swingby-3 rev/outbound Earth.4rev/inbound x 8 Fig. 7: Waiting orbit consisting of one-year and.4-year Earth revolution synchronous orbits. Drawn in the J inertial ecliptic coordinate system. ures 9 show example results of designed trajectories of the sub-payload in the case of the launch in 4. Black, blue and red lines indicate orbits of spacecraft, Earth and Mars, respectively. After the launch and tracing of a.4, or.5-year Earth revolution synchronous orbit, the sub-payload will execute the first Earth gravity assist. Second and third ones will be performed in the 6 winter and 8 spring. Aero-gravity assist can be conducted at either the first or second swingby. After the third gravity assist, the spacecraft will steer its course toward Mars and arrive in the winter of 8. The arrival velocity can be reduced more than km/s at maximum by means of preceding aero-gravity assists at Earth, although the reducible delta-v depends on capability of the thermal protection system for aeroassist. Figure also illustrates the same orbits in the -Earth line fixed rotational coordinate system. You can see combinations of characteristic patterns of synchronous orbits. 4. Trajectory Design Examples for Venus Earth rev/outbound.4rev/inbound In the same way as for Mars, we also designed a trajectory of the sub-payload for Venus. Because a good window for the Venus transfer exists in the winter of 6, we adopted an interplanetary parking orbit by using a -rev Earth revolution synchronous orbit between the winter of 5 and the winter of 6. Figure shows a result of the designed trajectory toward Venus to be launched in the winter of 4. Black, blue and red lines indicate orbits of spacecraft, Earth and Venus, respectively. Figure also illustrates the same orbit in the -Earth line fixed rotational coordinate system Fig. 8: Waiting orbit consisting of one-year and.4-year Earth revolution synchronous orbits. Drawn in the -Earth line fixed rotational coordinate system. 5 Summary This paper described a strategy to enhance mission capability for sub-payloads by incorporating Earth revolution syn- 4

5 x 8 x 8.5 4: Launch 5: Swingby- 6: Swingby-.5 Y [km] (J. Ecliptic) : Swingby-3 8: Arrival Y [km] (J. Ecliptic) X [km] (J. Ecliptic) x 8 Fig. 9: An example trajectory of the sub-payload to Mars launched in 4. Drawn in the J inertial ecliptic coordinate system. Black: Spacecraft, Blue: Earth, Red: Mars X [km] (J. Ecliptic) x 8 Fig. : An example trajectory of the sub-payload to Venus after the swingby in the winter of 5. Drawn in the J inertial ecliptic coordinate system. Black: Spacecraft, Blue: Earth, Red: Venus Fig. : An example trajectory of the sub-payload spacecraft launched in 4. Drawn in the -Earth line fixed rotational coordinate system Fig. : An example trajectory of the sub-payload to Venus after the swingby in the winter of 5. Drawn in the - Earth line fixed rotational coordinate system. 5

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