Flight dynamics results from the OEDIPUS-C tether mission

Tyc, George; Vigneron, F. R.; Jablonski, Alexander M.; Han, Ray P. S.; Modi, V. J.; Misra, A. K.

doi:10.2514/6.1996-3573

Cited by 12 publications

(4 citation statements)

References 5 publications

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“…1994). The mission was highly successful and the fight data indicated that the system dynamics behaved very much as expected (Tyc et al . 1996).…”

Section: Introductionsupporting

confidence: 62%

On the dynamics of spinning tethered space vehicles

Tyc¹,

Han

2001

Philosophical Transactions of the Royal Society of London. Seri

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Extensive work on the development of high-fidelity models of tethered systems has been carried out over the past three decades in support of the various space-tether missions flown to date. Although many of these models account for the full sixdegrees-of-freedom motions of the end bodies, they did not consider a spinning tether. The reason is that most of the investigations are believed to focus on specific mission configurations such as the shuttle-based and space station-based applications. They addressed the particular features for that class of tethered systems and a spinning tether was not one of those features. However, from the flight of Canadian OEDIPUS-A (Observations of Electric-field Distributions in an Ionospheric Plasma-a Unique Strategy) mission, which was the first to involve a spinning tether, it is clear that the spin of the tether is an important factor that must be included in the modelling. This is especially true for a wire-like tether which has a finite bending stiffness. In this paper, a set of nonlinear equations of motion is formulated for a spacecraft comprising two rigid payloads with flexible booms and connected by a tether that is spinning at the same nominal rate as the end bodies. Additionally, closed-form infinitesimal and quasi-stability conditions that provide very useful design equations for a spinning tethered spacecraft are presented. One of the main findings of this work is the dominant role played by the subtle effects of the tether root bending on the dynamics of a spinning tethered spacecraft, particularly at high spin rates. Using our model, we show that these effects lead to the dynamics anomaly observed in the OEDIPUS-A mission, when the aft payload exhibited an unexpectedly rapid and large divergence of the coning angle.

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“…1994). The mission was highly successful and the fight data indicated that the system dynamics behaved very much as expected (Tyc et al . 1996).…”

Section: Introductionsupporting

confidence: 62%

On the dynamics of spinning tethered space vehicles

Tyc¹,

Han

2001

Philosophical Transactions of the Royal Society of London. Seri

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“…Several interesting space applications of tethers have been proposed and several missions have been flown [Cosmo & Lorenzini 1997]; some missions were successful and others were unsuccessful. The major successful missions include the Canadian Space Agency's observation of Electric-Field Distributions on the Ionospheric Plasma-A Unique Strategy (OEDIPUS) missions: OEDIPUS-A in 1989 [Tyc & Han 1995] and OEDIPUS-C in 1995 [Tyc et al 1997], NASA's Small Expendable Deployer System (SEDS) missions: SEDS-1 in 1993 and SEDS-2 in 1994 [Smith 1995], NASA's Plasma Motor Generator (PMG) experiment in 1993 [Jost & Chlouber 1995], and the U.S. Naval Research Laboratory's tether physics and survivability (TiPS) in 1996 [Purdy et al 1997].…”

Section: Tethered Spacecraft Systemsmentioning

confidence: 99%

Fault Tolerant Control Of Spacecraft

Godard¹

2021

Preprint

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Autonomous multiple spacecraft formation flying missions demand the development of reliable control systems to ensure rapid, accurate, and effective response to various attitude and formation reconfiguration commands. Keeping in mind the complexities involved in the technology development to enable spacecraft formation flying, this thesis presents the development and validation of a fault tolerant control algorithm that augments the AOCS on-board a spacecraft to ensure that these challenging formation flying missions will fly successfully. Taking inspiration from the existing theory on nonlinear control, a fault-tolerant control system for the RyePicoSat missions is designed to cope with actuator faults whilst maintaining the desirable degree of overall stability and performance. Autonomous fault tolerant adaptive control scheme for spacecraft equipped with redundant actuators and robust control of spacecraft in underactuated configuration, represent the two central themes of this thesis. The developed algorithms are validated using a hardware-in-the-loop simulation. A reaction wheel testbed is used to validate the proposed fault tolerant attitude control scheme. A spacecraft formation flying experimental testbed is used to verify the performance of the proposed robust control scheme for underactuated spacecraft configurations. The proposed underactuated formation flying concept leads to more than 60% savings in fuel consumption when compared to a fully actuated spacecraft formation configuration. We also developed a novel attitude control methodology that requires only a single thruster to stabilize three axis attitude and angular velocity components of a spacecraft. Numerical simulations and hardware-in-the-loop experimental results along with rigorous analytical stability analysis shows that the proposed methodology will greatly enhance the reliability of the spacecraft, while allowing for potentially significant overall mission cost reduction.

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“…x k D e ®k t fU k cos º k t ¡ V k sin º k tg (12) Computer software has been written to solve the eigenvalueproblem to yield modal frequencies,modal convergence/divergence,i.e., growth/decay, ratios, and time-varying mode shapes, for given input parameters (inertias of the end bodies, spin rate, mass, stiffness damping, and length of the booms, and mass, damping tension, and length of the tether). Equation (12) is the basis for the computergenerated animation of the mode shapes.…”

Section: Mathematical Modelmentioning

confidence: 99%

Damped Gyroscopic Modes of Spinning Tethered Space Vehicles with Flexible Booms

Vigneron

Jablonski

Chandrashaker³

et al. 1997

Journal of Spacecraft and Rockets

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The OEDIPUS rocket payload con guration comprises two spinning subpayloads, each having long exible booms and a common long connecting tether of about 1 km in length. An analysis demonstrates the damped gyroscopic natural modes as a means of understanding the dynamics and stability of these very complex con gurations. For OEDIPUS-A, attention is focused on calculation of nutation mode divergence that is attributable to material damping of the tether and booms. Calculated divergence rates are compared with ight data of OEDIPUS-A. For OEDIPUS-C attention is focused on attitude and con guration stability. It is noted to be stable in nutation and other modes at its ight spin rate of about 0.09 Hz. However, the con guration is noted to be susceptible to structural instability at spin rates above 0.26 Hz. NomenclatureA; B; C = moments of inertia of end body, booms, and tether mass effects, kg-m 2 A; B= constant system matrices of rst-order form of model= distance between rotation point of end body and tether attachment (Fig. 2b), m C b1 ; C b2 = material linear viscous damping coef cient of booms C t = material linear viscous damping coef cient for tether D = symmetric damping matrix of con guration E I = stiffness of booms, N-m 2 E 1 ; E 2 = matrix functionals in stiffness matrices of booms G = skew-symmetric gyroscopic stiffness matrix K = stiffness matrix of con guration K T = stiffness submatrix of tether, N-m K 1 ; K 2 = stiffness submatrices of booms, N-m = length of tether, m 1 ;`2 = length of booms (Fig. 2c), m M = symmetric mass matrix of con guration M T = tether mass submatrices, kg-m 2 M 1 ; M 2 = boom mass submatrices, kg-m 2 m; n = order of discretization O = center of mass and of rotation of end body O X 1 X 2 X 3 = inertial coordinate system Ox 1 x 2 x 3 = coordinate system attached to rigid end body P = origin of tether reference system at attachment point of tether to end body when con guration is undeformed PY 1 Y 2 Y 3 = tether coordinate system that rotates with end body P 1 ; P 2 = column matrix functions of discretized booms [Eq. (6)], kg-m 2 Q k = complex eigenvector q 5 ; q 6 ; q 7 ; q 8 = coordinate function matrices of boom deformation q s 5 ; q s 6 ; q a 7 ; q a 8 = symmetric and antisymmetric coordinate function matrices of booms R = column matrix functionals of the tether [Eq. (6)], kg-m 2 S = spin rate, cps s = eld point of tether or boom length, m T = tether tension, N t = time, s U k ; V k = real and imaginary parts of Q k u; u; v; w = tether deformation vector and components (Fig. 2a), m w i = boom de ection variables x = state variable matrix for model in second-order form Z = state variable matrix for model in rst-order form ® k ; º k = real and imaginary part of eigenvalue, 1/s ® N ; ® t = real part of eigenvalue of nutation mode and tether modes, respectively, 1/s = material damping ratios of tether or booms µ ; Ã; Á = Euler angles of lower end body (Fig. 2b), deģ k = complex eigenvalue » = dimensionless length of tether or booms ½ ; ½ 1 ; ½ 2 = density per unit length of tether an...

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Flight dynamics results from the OEDIPUS-C tether mission

Cited by 12 publications

References 5 publications

On the dynamics of spinning tethered space vehicles

On the dynamics of spinning tethered space vehicles

Fault Tolerant Control Of Spacecraft

Damped Gyroscopic Modes of Spinning Tethered Space Vehicles with Flexible Booms

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