For medium and low latitudes, the N-hour repeating equatorial orbit is a new option for remote sensing satellites, which can realize multiple revisits per day, and revisit at the same local time revisits on different days. The remote sensing constellation formed by equatorial satellites could easily achieve rapid revisit, or even continuous observation. Therefore, the ground surveillance performance is critical for the system design of equatorial remote sensing constellations. In this paper, the latitude coverage characteristics of equatorial satellites are analyzed. This paper also assesses the constellation size required to achieve continuous observation of targets in different latitudes. The staring imaging time durations of satellites in the prograde and retrograde constellation are analyzed respectively.
Sun-synchronous orbit (SSO) has a characteristic that passes the given target area at the same local mean solar time, and is commonly used by remote sensing satellites. The ground track of these satellites could repeat precisely, by applying specific design methods. However, the revisit time of a single SSO remote sensing satellite could be greater than one day, and hard to achieve continuous Earth observation, especially for low latitudes. The N-hour repeating equatorial orbit introduced in this paper is a new concept medium Earth orbit (MEO) for remote sensing satellites, which can realize several revisits in one day, and same local time revisits on different days for a designated target at low latitudes. In this paper, the inertial period definition of equatorial satellites is clearly defined and calculated, the design of prograde and retrograde repeating equatorial orbits is given and analyzed theoretically and numerically. The longitude coverage features of equatorial satellites are also evaluated.
In order to formulate accurate and reasonable random vibration test conditions for optical cameras and solve the problem of conservative design of test conditions caused by inaccurate simulation calculation, this paper proposes a general method for the design of random vibration specification for small satellites’ optical cameras. First, we derive the dynamic response formula, which lays a foundation for the small satellite dynamic simulation calculation. Then, the response of the optical camera mounting surface, which is obtained by the dynamic calculation of the small satellite finite element model, and the random vibration test conditions of the camera process is enveloped to obtain the preliminary random vibration test condition of the new research optical camera. Subsequently, we combine the random vibration test data of the optical remote sensing satellite, which has been used in orbit, and revise the preliminary random vibration test conditions of the new research camera in terms of response magnitude and frequency. Finally, we verify the rationality of this method through random vibration test of a small satellite. This scheme is conducive to the establishment of more reasonable and feasible random vibration test conditions for small satellites’ optical cameras, which is beneficial to the development and production of optical cameras.
Equatorial satellites in low or medium altitude orbits have shown great potential for monitoring space debris in the geosynchronous belt and performing remote sensing tasks in low latitudes. However, the possibility of satellite collision between proposed equatorial satellites and existed spacecraft in Sun-synchronous orbit (SSO) should be considered. In this paper, the definition and design methods of both prograde and retrograde N-hour repeating equatorial orbits are given for space debris monitoring and remote sensing. The characteristics of prograde and retrograde equatorial orbits are compared. For reducing the risk of satellite collision after retirement, this paper also makes suggestions on the system design of the equatorial satellite.
The application of small satellites is a new focus on deep space exploration. Small satellites are easy to launch as secondary payloads, and more suitable for international cooperation. However, small satellites have several major weaknesses for deep space missions, such as the lack of orbit transfer capability and large antenna. This paper presents an innovative concept of small satellite bus based on Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA). The proposed satellite bus is equipped with orbit control subsystem and a large parabolic antenna for the exploration of Jovian and Saturnian systems. In this paper, the mission requirements, the launch feasibilities and the design parameters of the proposed satellite bus are also discussed.