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Supercapacitors, also known as ultracapacitors or Electric Double-Layer Capacitors, are considered as new solutions for space power supplies because of their advantages such as a high power-density, a wide temperature range and a long life. Graphene supercapacitors (GSCs), with graphene structure served as the highly active material, are new types of high-performance and new-concept supercapacitors. However, there are no cases of supercapacitors, especially GSCs, being used as power system for spacecraft independently. This paper introduces the design, test and application of the power supply based on GSCs for the electromagnetic separation system in Q-SAT mission. The power supply is integrated by multiple GSCs to meet the requirements of the separation system, providing 600W power within 2 × 10−5 m3 and 1 kg. The design, packaging method, test and result is reported in the paper. The rated voltage of the selected GSCs is 2.7 V, and the power density reaches 1.1 × 103 W/kg and 2.6 × 106 W/m3. The power supply based on GSCs has passed the verification test successfully including mechanical, electrical, vacuum and thermal tests from −20–60 °C. It successfully drove the separation system and guaranteed the task of unlocking and releasing Q-SAT in the flight on Aug 6, 2020. The significance of this work lies in the following aspects: First, the feasibility of application on the satellite separation system in space of GSC is demonstrated. Second, the design and application case of supercapacitor as a driving power supply for a satellite separation system is presented. Third, a screening and integration method for GSCs packaged as a space power supply is proposed and verified. The space power supply based on GSCs will be used more widely in space according to this research.

The population of nano satellites has increased dramatically in recent years, but there are few launch vehicles providing dedicated launch opportunities for them. This paper presents the conceptual design of an innovative nano air launch vehicle, dedicated to providing exclusive launch opportunity to nano satellites weighing 10 kg or less. This nano launcher is a three-stage launch vehicle carried by a twin-fuselage unmanned aerial vehicle and released at an altitude of 10 km from a velocity of 0.5Ma. The air launch vehicle is innovative in two senses: firstly, it adopts a novel air launch method. Before separation, the launch vehicle is adjusted to a large angle of attack by the carrier aircraft. The launch vehicle only needs to keep attitude stable after separation, and sequentially when first stage ignites, the orientation of the thrust is inclined upward so that the thrust can provide enough vertical force to pull up the flight path angle rapidly, without aid from large wings; secondly, a terminal lumped control unit is applied to undertake the flight control task on its own. The terminal lumped control unit is a combination of four thrusters that share the liquid propellant with the last stage's main engine. With the terminal lumped control unit, the air launch vehicle needs no movable attitude control instruments, including thrust vectoring gimbals and movable wings. This innovative flight control strategy makes the launch vehicle very simple in structure and light in weight. Multidisciplinary modeling of the nano launcher is performed and genetic algorithm is applied for optimization. Among all the disciplines, trajectory optimization is key in checking the feasibility of the design, for which the Radau pseudospectral method is successfully implemented to obtain the optimal ascent trajectory. The optimal nano launcher has a gross take-off weight of 1365 kg. Finally, a sensitivity analysis is performed to check the robustness of the result using Monte Carlo simulation method.

The self-organizing control strategy for asteroid intelligent detection swarm, which is considered as a space application instance of intelligent swarm, is developed. The leader-follower model for the asteroid intelligent detection swarm is established, and the further analysis is conducted for massive asteroid and small asteroid. For a massive asteroid, the leader spacecraft flies under the gravity field of the asteroid. For a small asteroid, the asteroid gravity is negligible, and a trajectory planning method is proposed based on elliptic cavity virtual potential field. The self-organizing control strategy for the follower spacecraft is developed based on a mechanism of velocity planning and velocity tracking. The simulation results show that the self-organizing control strategy is valid for both massive asteroid and small asteroid, and the exploration swarm forms a stable configuration.

The responsive orbits play a key role in addressing the mission of Operationally Responsive Space (ORS) because of their capabilities. These capabilities are usually focused on supporting specific targets as opposed to providing global coverage. One subtype of responsive orbits is repeat coverage orbit which is nearly circular in most remote sensing applications. This paper deals with a special kind of repeating ground track orbit, referred to as circular revisit orbit. Different from traditional repeat coverage orbits, a satellite on circular revisit orbit can visit a target site at both the ascending and descending stages in one revisit cycle. This typology of trajectory allows a halving of the traditional revisit time and does a favor to get useful information for responsive applications. However the previous reported numerical methods in some references often cost lots of computation or fail to obtain such orbits. To overcome this difficulty, an analytical method to determine the existence conditions of the solutions to revisit orbits is presented in this paper. To this end, the mathematical model of circular revisit orbit is established under the central gravity model and the J2 perturbation. A constraint function of the circular revisit orbit is introduced, and the monotonicity of that function has been studied. The existent conditions and the number of such orbits are naturally worked out. Taking the launch cost into consideration, optimal design model of circular revisit orbit is established to achieve a best orbit which visits a target twice a day in the morning and in the afternoon respectively for several days. The result shows that it is effective to apply circular revisit orbits in responsive application such as reconnoiter of natural disaster.

To meet the very demanding requirements for space gravity detection, the gravitational reference sensor (GRS) as the key payload needs to offer the relative position of the proof mass with extraordinarily high precision and low disturbance. The position determination and error analysis for the GRS with a spherical proof mass is addressed. Firstly the concept of measuring the freely falling proof mass with optical shadow sensors is presented. Then, based on the optical signal model, the general formula for position determination is derived. Two types of measurement system are proposed, for which the analytical solution to the three-dimensional position can be attained. Thirdly, with the assumption of Gaussian beams, the error propagation models for the variation of spot size and optical power, the effect of beam divergence, the chattering of beam center, and the deviation of beam direction are given respectively. Finally, the numerical simulations taken into account of the model uncertainty of beam divergence, spherical edge and beam diffraction are carried out to validate the performance of the error propagation models. The results show that these models can be used to estimate the effect of error source with an acceptable accuracy which is better than 20%. Moreover, the simulation for the three-dimensional position determination with one of the proposed measurement system shows that the position error is just comparable to the error of the output of each sensor.

The Inner Formation Flying System (IFFS) consisting of an outer satellite and an inner satellite which is a solid sphere proof mass freely flying in the shield cavity can construct a pure gravity orbit to precisely measure the earth gravity field. The gravitational attraction on the inner satellite due to the outer satellite is a significant disturbance source to the pure gravity orbit and is required to be limited to 10 −11 m s −2 order. However, the gravitational disturbance force was on 10 −9 m s −2 order actually and must be reduced by dedicated compensation mass blocks. The region of relative motion of the inner satellite about its nominal position is within 1 cm in dimension, which raises the complexity of the compensation blocks design. The iterative design strategy of the compensation blocks based on reducing the gravitational attraction at the nominal position of the inner satellite is presented, aiming to guarantee the gravitational force in the relative motion region within requirements after the compensation. The compensation blocks are designed according to the current status of IFFS, and the gravitational disturbance force in the region is reduced to 10 −11 ms −2 order with minimized adding mass.

Inner-Formation Gravity Measurement Satellite System (IFGMSS) is used to explore the Earth gravity using two satellites in an inner-formation flying mode. To fulfill the mission, relative position of the two satellites is required to be zeroed and attitude of the outer satellite should be stabilized in real-time. This paper proposed an integrated control scheme for the IFGMSS, and the main idea is to use only thrusters to control the relative position and attitude. The integrated control loop contains a control law's module and a control allocation law's module. The control law based on the feedback linearization method makes nonlinear dynamics counteracted and uses PD control law to reformulate the dynamics into a linear form. The integrated control allocation law is designed to assign the commanded control force and moment to each thruster dynamically. We transfer the control allocation problem into a linear programming (LP) problem and use the Optimal Theory to calculate the corresponding thrust of each thruster. Finally, an IFGMSS mission is simulated, where the two satellites fly in a circle orbit with a 300 km's altitude. Results using the integrated control scheme and the traditional separated control scheme are compared and analyzed. It has been found that the integrated control scheme is superior to the separated control scheme in output ability, level of redundancy and fuel cost.

A spherical proof mass floating in a cage of an inner-formation satellite sometimes needs to change its orbital elements. This occurs in a case that the position of the proof mass is not in the target orbit. To protect the proof mass from the mechanical damage, a scheme of solar radiation is proposed. The solar beam is collected by a lens system and projected on the proof mass when commanded. However the solar radiation may change the temperature of the proof mass. So this paper further investigates the heat radiation issue between the proof mass and the inner wall of the cage. It has been found that the heat balance can be achieved by selecting proper parameters (such as radiant emittance and absorbency) of the proof mass and the cage, so that the temperature of the proof mass could be stabilized. The numerical results show that the solar radiation can produce a force with a magnitude of μN in the condition of steady heat balance.

Formation keeping is a key technology of distributed satellite system. In order to deal with the satellite formation keeping problem including perturbations, normalized linear equations of satellite relative motion including J 2 perturbation are derived. Then a sliding mode variable structure control with matching uncertainty and disturbance is introduced, the method using a reaching law to design VSC controller directly for the discrete time system. Then control law for formation keeping of satellite formation in LEO with small eccentricity reference orbit is designed, so the modeling error and disturbance is included. A simulation environment with layer control architecture and nonlinear dynamics including perturbations are developed to verify the control algorithms. Simulations for various system parameters are conducted for both circular reference orbit and eccentric reference orbit. Simulation results established the robustness and performance of the controller.