Lunar science with affordable small spacecraft technologies: MoonLITE and Moonraker

Gao, Yang; Phipps, Andy; Taylor, M.; Crawford, Ian; Ball, Andrew; Wilson, Lionel; Parker, Dave; Sweeting, Martin; Curiel, Alex da Silva; Davies, P.G.; Baker, Adam; Pike, Tom; Smith, A.; Gowen, R. A.

doi:10.1016/j.pss.2007.11.005

Cited by 30 publications

(13 citation statements)

References 15 publications

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“…The angle calculated from the simulation results was found to be less than five degrees with an angle of attack of four degrees calculated from the remaining attitude error; both angles were less than previous mission criteria [7], [13]. The impact error ellipse is calculated from residual velocities causing drift and was found to be 8.3km x 6.8km, which was again within mission stated objectives [4], [13]. …”

Section: Table1 Lunar Penetrator Mass and Inertia Values Used For Simentioning

confidence: 57%

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Attitude Control of Small Probes for De-Orbit, Descent and Surface Impact on Airless Bodies Using a Single PWM Thruster

Gillham¹,

Howells²

2017

2017 6th International Conference on Space Mission Challenges for Information Technology (SMC-IT)

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Abstract-A single thruster attitude and de-orbital control method is proposed, capable of delivering a small spin stabilized probe with payload to the surface of an airless body such as the Moon. Nutation removal, attitude control and fast large angle maneuvers have been demonstrated and shown to be effective using a commercially available single standard cold gas pulse width modulated controlled thruster model. Maximum final impact angle due to drift and residual velocities was found to be less than 5 degrees and the maximum angle of attack to be 4 deg. The conventional 3-axis control would require as many as twelve thrusters requiring a more substantial structure with complex pipework, and a more sophisticated controller. The single thruster concept minimises the mass requirement and thus cost of the mission, making the concept of small networked surface probes for extended science missions more viable. Experiments based on computer simulation have shown that strict design and mission profile requirements can be fulfilled using the single thruster control method.

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Section: Table1 Lunar Penetrator Mass and Inertia Values Used For Simentioning

confidence: 57%

“…The UK MoonLITE mission proposal's main goal was to be seismic detection hence investigating the lunar internal structure. Infra-red spectroscopic sensors capable of volatile substance analysis, which would be invaluable data for any long term lunar manned presence [4].…”

Section: Introductionmentioning

confidence: 99%

Attitude Control of Small Probes for De-Orbit, Descent and Surface Impact on Airless Bodies Using a Single PWM Thruster

Gillham¹,

Howells²

2017

2017 6th International Conference on Space Mission Challenges for Information Technology (SMC-IT)

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“…In this paper, Sliding Mode Control (SMC) is chosen as the basic control method for spacecraft's attitude control due to its advantages especially for small spacecraft space exploration, such as LunarSat [5]. Specifically, SMC is well-known as a robust controller, it is low complexity, can have low computational burden, low weight and low cost [6] [7].…”

Section: Introductionmentioning

confidence: 99%

Application of decaying boundary layer and switching function method thorough error feedback for sliding mode control on spacecraft's attitude

Hassrizal

Rossiter

2017

2017 25th Mediterranean Conference on Control and Automation (MED)

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Abstract-Effective operation of small spacecraft implies processors with low cost, energy efficiency and low computational burdens while retaining accurate output tracking. This paper presents the extension of work in [1] on eliminating the chattering for Sliding Mode Control (SMC) using a decaying boundary layer design which is able to achieve these small spacecraft operation needs. The extension is applied on a spacecraft's attitude control, while orbiting the earth with angular velocity, ω0. In SMC, chattering is a main drawback as it can cause wear and tear to moving mechanical parts. Earlier work on a decaying boundary layer design was capable of reducing the chattering phenomena for a limited time only and hence this paper proposes a novel decaying boundary layer and switching function to improve the earlier version. The proposed technique is shown to reduce chattering permanently and also retain control output accuracy.

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“…Pulse-train algorithms can also be used for oblate spacecraft. The existing slew algorithms have been initially developed for specific prolate spacecraft such as penetrators proposed in MoonLITE missions [6]. In the MoonLITE mission, a mothership releases missile-shaped penetrators equipped with thrusters for hard landing on the lunar surface from 100 km altitude.…”

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confidence: 99%

Slew Control of Prolate Spinners Using Single Magnetorquer

Gao

Chanik

2016

Journal of Guidance, Control, and Dynamics

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= angular momentum of spacecraft, that in reference inertial/spacecraft-fixed body coordinates, that after H t = the first actuation, that at end of the slew, and the target angular momentum, N · ms I, I z , I t = spacecraft moment of inertia matrix, moment of inertia of transverse axis/spin axis, kg · m 2 k = rotation number around spin axis of the spacecraft m = magnetic moment of the magnetorquer dipole, A · m 2 T, T magnet = control torque/ magnetic torque applied to the spacecraft, N · m t 0 , t 1f , t 2s , t 2f , t 1;2 , Δt 2 , t ⌢ 2 = time of the start, time at the end of the first actuation, time of the start of the second actuation, time at the end of second actuation, time interval between first and second actuation, the actuating duration of the second actuation and the predicted meantime of second actuation, s Z 0 , Z 1f , Z f , Z ins , Z t = spin axis of the initial/after the first actuation/after the slew/of an instant time and of the target orientation β = slew angle, deg γ = residual precession angle after the first actuation of slew, deg θ = nutation angle, deg λ = inertia ratio of spin axis and transverse axis moment of inertia ξ = the second actuation time adjustment coefficient ω, ω Z , ω XY , ω N , ω H = angular velocity of the spacecraft, that in spin axis, that in the X − Y plane, body nutation rate, and inertial nutation rate, rad∕s

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Lunar science with affordable small spacecraft technologies: MoonLITE and Moonraker

Cited by 30 publications

References 15 publications

Attitude Control of Small Probes for De-Orbit, Descent and Surface Impact on Airless Bodies Using a Single PWM Thruster

Attitude Control of Small Probes for De-Orbit, Descent and Surface Impact on Airless Bodies Using a Single PWM Thruster

Application of decaying boundary layer and switching function method thorough error feedback for sliding mode control on spacecraft's attitude

Slew Control of Prolate Spinners Using Single Magnetorquer

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