On-orbit Mass Research Articles

Mission architecture analysis for manned NEA exploration using MAM method

Near Earth Asteroids (NEAs) have become a hot topic these years for their scientific interests and technological accessibility. Manned NEA mission is a good choice for asteroid on-site exploration as well as a stepping stone for future manned deep space missions. Focusing on manned NEAs mission architecture analyzing, this paper uses the mission architecture matrix (MAM) method to reduce mission mass scale, for available studies show that simple mission architectures achieving NEAs all require LEO launching mass exceeding available space carrying capacity. By developing a series of mission architectures based on manned spacecraft system design method, the mass scale is estimated according to mission data of 63 most accessible NEA targets. It is discovered that with advanced mission architectures, maximum LEO launching mass can be significantly reduced by separate launching, rendezvous, and docking. The original on-orbit mass about 300–400 metric tons (mT) is reduced to 200–250 mTs, and the maximum launching mass is reduced to only about 70–100 mTs in several cases. This provides us a high possibility to launch manned NEA missions in the near future. In addition, this method can be used in mission mass scale analysis to other deep space targets. Advanced mission architecture also brings challenges on reliability, which plays another key role in mission design. The balance between mass scale and reliability needs to be kept in future works.

Target selection and mass estimation for manned NEO exploration using a baseline mission design

In recent years Near-Earth Objects (NEOs) have received an increased amount of interest as a target for human exploration. NEOs offer scientifically interesting targets, and at the same time function as a stepping stone for achieving future Mars missions. The aim of this research is to identify promising targets from the large number of known NEOs that qualify for a manned sample-return mission with a maximum duration of one year. By developing a baseline mission design and a mass estimation model, mission opportunities are evaluated based on on-orbit mass requirements, safety considerations, and the properties of the potential targets. A selection of promising NEOs is presented and the effects of mission requirements and restrictions are discussed. Regarding safety aspects, the use of free-return trajectories provides the lowest on-orbit mass, when compared to an alternative design that uses system redundancies to ensure return of the spacecraft to Earth. It is discovered that, although a number of targets are accessible within the analysed time frame, no NEO offers both easy access and high incentive for its exploration. Under the discussed aspects a first human exploration mission going beyond the vicinity of Earth will require a trade off between targets that provide easy access and those that are of scientific interest. This lack of optimal mission opportunities can be seen in the small number of only 4 NEOs that meet all requirements for a sample-return mission and remain below an on-orbit mass of 500 metric Tons (mT). All of them require a mass between 315 and 492mT. Even less ideal, smaller asteroids that are better accessible require an on-orbit mass that exceeds the launch capability of future heavy lift vehicles (HLV) such as SLS by at least 30mT. These mass requirements show that additional efforts are necessary to increase the number of available targets and reduce on-orbit mass requirements through advanced mission architectures. The need for on-orbit assembly also becomes apparent, as availability of a HLV alone does not provide sufficient payload capabilities for any manned mission targeting NEOs.

New directions for space solar power

Several of the central issues associated with the eventual realization of the vision of solar power from space for terrestrial markets resolve around the expect costs associated with the assembly, inspection, maintenance and repair of future solar power satellite (SPS) stations. In past studies (for example, NASA's “Fresh Look Study”, c. 1995–1997) efforts were made to reduce both the scale and mass of large, systems-level interfaces (e.g., the power management and distribution (PMAD) system) and on-orbit fixed infrastructures through the use of modular systems strategies. These efforts have had mixed success (as reflected in the projected on-orbit mass of various systems concepts. However, the author remains convinced of the importance of modular strategies for exceptionally large space systems in eventually realizing the vision of power from space. This paper will introduce some of the key issues associated with cost-competitive space solar power in terrestrial markets. It will examine some of the relevant SPS concepts and will assess the ‘pros and cons’ of each in terms of space assembly, maintenance and servicing (SAMS) requirements. The paper discusses at a high level some relevant concepts and technologies that may play r role in the eventual, successful resolution of these challenges. The paper concludes with an example of the kind of novel architectural approach for space solar power that is needed.

HF Satellite Sound Broadcasting - A Preliminary Assessment

This paper examines in a preliminary way the technical factors associated with the use of the HF band for satellite sound broadcasting. Numerical results are presented for an example geostationary satellite system operating in the 26 MHz band. An analysis of the electrical characteristics and the far-field pattern of a whip antenna used on low-cost HF portable receivers is presented in order to connect receiver noise-limited sensitivity measurements and signal quality objectives with incident field strength requirements. It is shown that a minimum usable circularly polarized field strength of 390 μV/m is required for an elevation angle of 5° in a noisy urban environment, decreasing to about 130 μV/m above 15° elevation in a quiet rural area. A simplified satellite mass model is used to estimate the minimum mass of a satellite to provide a field strength of 1000 μV/m at the edge of the coverage area. The optimum antenna size is determined as a function of the antenna parametrics, e.i.r.p., and frequency. For the conditions assumed, the spacecraft on-orbit mass will be on the order of 14 000-17 000 kg, and the antenna diameter will be about 600 m. The Appendix contains analyses showing that refraction, scintillation, and absorption caused by transionospheric propagation are the principal mechanisms degrading the received signal quality. An analysis using ray tracing techniques shows that refraction will reduce coverage to about 40 percent of the earth's disk during the daytime for periods when the smoothed sunspot number is 100. Nighttime coverage will; because of refraction, extend beyond the visible horizon. Nighttime scintillation in the equatorial region is expected to be severe. Based on an extrapolation of data taken at higher frequencies, the envelope of the received carrier will exhibit Rayleigh statistics with fade depths on the order of 10 dB occurring on the average of one every 10 s and lasting for about 1 s. Absorption of transionospheric signals occurs primarily in the <tex xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">D</tex> region of the ionosphere and amounts to a nominal 2 dB of loss. During periods of sudden ionospheric disturbances due to solar flares the absorptive losses may approach 40 dB.