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Escape Maneuvers with Lunar Gravity Assists

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Escape Maneuvers with Lunar Gravity Assists POLITECNICO DI TORINO Facolt`adi Ingegneria Corso di Laurea in Ingegneria Aerospaziale Tesi di Laurea Magistrale Escape maneuvers with lunar gravity assists Relatore: prof. Lorenzo Casalino Candidato: Alex Siniscalco Anno accademico 2017-2018 ‘Nothing is achieved without sacrifice and courage . Who is capable to suffer will eventually win’ Mahatma Gandhi Abstract For spacecraft interplanetary missions, gravity assist maneuvers are usually used because they allow one to get as a rule high ∆V without having to use propellant, but exploiting the energy given by the gravitational attraction of a celestial body. This allows improving the useful mass for a specific mission, for a modest increase in flight time. Thus, when escape from Earth’s sphere of influence is regarded, lunar gravity assist maneuvers become a valid alternative to direct escapes. In this paper, two consecutive gravity assist maneuvers are considered and the possible moon-to-moon trajectories that allow them are sought. Afterwards, considering different spacecraft initial velocities relative to Moon, the effect on the escape orbit is analyzed, with the purpose to link the features of the latter (i.e the heliocentric flight-path angle and declination) to the maximum obtainable specific mechanical energy (i.e. the achievable escape velocity). First, a brief introduction will be presented, followed by a chapter dedicated to the mathematical models adopted to carry out the thesis; after that, will be illustrated the technique used for the study of moon-to-moon orbits and lunar gravity assists, followed by a description of the attained results. Finally, the conclusive observations, in addition to the problems faced during the analysis and a short debate on the possible future developments of the present work, will be introduced in a dedicated chapter. Contents 1 Introduction1 2 Mathematical model5 2.1 The N-body problem.......................5 2.2 The 2-body problem....................... 10 2.3 Constants of the motion..................... 12 2.3.1 Specific angular momentum............... 12 2.3.2 Specific mechanical energy................ 13 2.4 Equation of the trajectory.................... 15 2.5 The circular restricted 3-body problem............. 16 2.6 Gravity assist........................... 18 2.6.1 The patched-conic approximation............ 18 2.6.2 Lunar gravity assist.................... 19 3 Analysis 23 3.1 Solar-perturbed moon-to-moon transfer............. 23 3.2 Moon-to-escape lunar gravity assist............... 27 3.2.1 Lunar gravity assist evaluation............. 27 3.2.2 Maximum C3, given heliocentric flight-path angle and declination......................... 34 4 Results 35 4.1 Moon-to-moon leg......................... 35 4.1.1 Outbound-inbound.................... 35 4.1.2 Inbound-inbound..................... 40 I II Contents 4.2 Lunar gravity assist........................ 47 4.2.1 C3(γ) for given δ ..................... 47 4.2.2 Global results....................... 71 5 Conclusions 77 5.1 Conclusions............................ 77 5.2 Future work............................ 78 List of Tables ◦ 4.1 Ranges (in ) for α and θSC for oi families and V1 = 0:9 Km/s. 39 ◦ 4.2 Ranges (in ) for α and θSC for oi families and V1 = 1 Km/s. 39 ◦ 4.3 Ranges (in ) for α and θSC for oi families and V1 = 1:1 Km/s. 39 ◦ 4.4 Ranges (in ) for α and θSC for ii families and V1 = 0:9 Km/s. 44 ◦ 4.5 Ranges (in ) for α and θSC for ii families and V1 = 1 Km/s.. 44 ◦ 4.6 Ranges (in ) for α and θSC for ii families and V1 = 1:1 Km/s. 44 III List of Figures 2.1 The N-body problem.......................8 2.2 Relative motion of 2 bodies.................... 11 2.3 Zenit and flight-path angle.................... 13 2.4 The 3-body problem in an inertial reference........... 17 2.5 Lunar gravity assist manoeuvre schema............. 21 2.6 Components of V1 expressed via pump (p) and crank (k) angles. 22 3.1 Problem schema.......................... 24 3.2 Geometrical schema for escape orbit............... 32 3.3 Heliocentric flight-path angle (γH ) and declination (δH ) for the spacecraft’s velocity after the lunar gravity assist........ 33 4.1 Outbound-inbound orbits for V1 = 0:9 Km/s.......... 36 4.2 Outbound-inbound orbits for V1 = 1 Km/s........... 37 4.3 Outbound-inbound orbits for V1 = 1:1 Km/s.......... 38 4.4 Inbound-inbound orbits for V1 = 0:9 Km/s........... 41 4.5 Inbound-inbound orbits for V1 = 1 Km/s............ 42 4.6 Inbound-inbound orbits for V1 = 1:1 Km/s........... 43 4.7 Range of α that allow a rendezvous for oi and ii families.... 45 4.8 Range of θ that allow a rendezvous for oi and ii families.... 46 4.9 C3(γ) for fixed δ and V1 = 0:9 Km/s, Aoi family....... 48 4.10 C3(γ) for fixed δ and V1 = 1 Km/s, Aoi family........ 48 4.11 C3(γ) for fixed δ and V1 = 1:1 Km/s, Aoi family....... 49 4.12 C3(γ) for fixed δ and V1 = 0:9 Km/s, Aii family....... 49 4.13 C3(γ) for fixed δ and V1 = 1 Km/s, Aii family........ 50 V VI List of Figures 4.14 C3(γ) for fixed δ and V1 = 1:1 Km/s, Aii family....... 50 4.15 C3(γ) for fixed δ and V1 = 0:9 Km/s, Boi family....... 52 4.16 C3(γ) for fixed δ and V1 = 1 Km/s, Boi family........ 52 4.17 C3(γ) for fixed δ and V1 = 1:1 Km/s, Boi family....... 53 4.18 C3(γ) for fixed δ and V1 = 0:9 Km/s, Bii family....... 53 4.19 C3(γ) for fixed δ and V1 = 1 Km/s, Bii family........ 54 4.20 C3(γ) for fixed δ and V1 = 1:1 Km/s, Bii family....... 54 4.21 C3(γ) for fixed δ and V1 = 0:9 Km/s, Coi family....... 56 4.22 C3(γ) for fixed δ and V1 = 1 Km/s, Coi family........ 56 4.23 C3(γ) for fixed δ and V1 = 1:1 Km/s, Coi family....... 57 4.24 C3(γ) for fixed δ and V1 = 0:9 Km/s, Cii family....... 57 4.25 C3(γ) for fixed δ and V1 = 1 Km/s, Cii family........ 58 4.26 C3(γ) for fixed δ and V1 = 1:1 Km/s, Cii family....... 58 4.27 C3(γ) for fixed δ and V1 = 0:9 Km/s, Doi family....... 60 4.28 C3(γ) for fixed δ and V1 = 1 Km/s, Doi family........ 60 4.29 C3(γ) for fixed δ and V1 = 1:1 Km/s, Doi family....... 61 4.30 C3(γ) for fixed δ and V1 = 0:9 Km/s, Dii family....... 61 4.31 C3(γ) for fixed δ and V1 = 1 Km/s, Dii family........ 62 4.32 C3(γ) for fixed δ and V1 = 1:1 Km/s, Dii family....... 62 4.33 C3(γ) for fixed δ and V1 = 0:9 Km/s, Eoi family....... 64 4.34 C3(γ) for fixed δ and V1 = 1 Km/s, Eoi family........ 64 4.35 C3(γ) for fixed δ and V1 = 1:1 Km/s, Eoi family....... 65 4.36 C3(γ) for fixed δ and V1 = 0:9 Km/s, Eii family....... 65 4.37 C3(γ) for fixed δ and V1 = 1 Km/s, Eii family........ 66 4.38 C3(γ) for fixed δ and V1 = 1:1 Km/s, Eii family....... 66 4.39 C3(γ) for fixed δ and V1 = 0:9 Km/s, F oi family....... 68 4.40 C3(γ) for fixed δ and V1 = 1 Km/s, F oi family........ 68 4.41 C3(γ) for fixed δ and V1 = 1:1 Km/s, F oi family....... 69 4.42 C3(γ) for fixed δ and V1 = 0:9 Km/s, F ii family....... 69 4.43 C3(γ) for fixed δ and V1 = 1 Km/s, F ii family........ 70 4.44 C3(γ) for fixed δ and V1 = 1:1 Km/s, F ii family....... 70 4.45 C3max(γ) for fixed δ and V1 = 0:9 Km/s............. 72 4.46 C3max(γ) for fixed δ and V1 = 1 Km/s.............. 72 List of Figures VII 4.47 C3max(γ) for fixed δ and V1 = 1:1 Km/s............. 73 4.48 Minimum C3max(γ) varying with δ and V1........... 74 4.49 Minimum Vescape varying with δ and V1............. 74 4.50 Minimum Vescape=V1 varying with δ and V1........... 75 4.51 Maximum C3max(γ) varying with δ and V1........... 75 4.52 Maximum Vescape varying with δ and V1............. 76 4.53 Maximum Vescape=V1 varying with δ and V1........... 76 Chapter 1 Introduction In order to develop an interplanetary mission there are two important as- pects to consider: the definition of the escape trajectory, which allows to the spacecraft to go out of the Earth’s sphere of influence (i.e. the spheroidal re- gion of space where the Earth gravitational pull prevails) and the associated maneuver, which allows to the spacecraft to reach the wanted heliocentric orbit. The choice of escape strategy used is important, since it affects the cost of the mission in reference to propellant necessary that directly influence the payload mass that can be carried. There are different escape strategies for a specific mission, those usually used are: • direct escape – a single-impulse maneuver leading the spacecraft on an hyperbolic orbit, that after is travelled until when the escape condi- tions are not reached, i.e. the wanted hyperbolic excess velocity; this strategy needs of a high-thrust propulsion system, that uses chemical propellants; • Oberth maneuver (powered flyby) - two single-impulse maneuvers are used that allow reducing the gravitational losses: the first allows de- creasing the spacecraft’s velocity while to move closer to the main body with an Hohmann transfer and the second, given to the periastron now lowered, bringing the spacecraft on an hyperbolic escape orbit, anal- ogous to the one of the direct escape.
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