Earth-to-orbit Research Articles

Biological challenges of true space settlement

“Space Settlements” – i.e., permanent human communities beyond Earth's biosphere – have been discussed within the space advocacy community since the 1970s. Now, with the end of the International Space Station (ISS) program fast approaching (planned for 2024–2025) and the advent of low cost Earth-to-orbit (ETO) transportation in the near future, the concept is coming once more into mainstream. Considerable attention has been focused on various issues associated with the engineering and human health considerations of space settlement such as artificial gravity and radiation shielding. However, relatively little attention has been given to the biological implications of a self-sufficient space settlement. Three fundamental questions are explored in this paper: (1) what are the biological “foundations” of truly self-sufficient space settlements in the foreseeable future, (2) what is the minimum scale for such self-sustaining human settlements, and (3) what are the integrated biologically-driven system requirements for such settlements? The paper examines briefly the implications of the answers to these questions in relevant potential settings (including free space, the Moon and Mars). Finally, this paper suggests relevant directions for future research and development in order for such space settlements to become viable in the future.

Concerning the Design of a Novel Electromagnetic Launcher for Earth-to-Orbit Micro- and Nanosatellite Systems

This paper presents an alternative launcher design for Earth-to-orbit (ETO) micro- and nanosatellite systems. The main goal of the design is to reduce the cost of the launching operation. This paper also addresses the calculation of the energy required for the launching operation and the selection of a power source for the designed satellite launcher. The method introduced here proposes a vertical-takeoff electromagnetic launcher for ETO micro- and nanosatellite systems. The design specifications are 7-km/s muzzle velocity and 30000-gee acceleration for a 10-kg payload. It is expected that the proposed design would present a viable alternative to conventional launchers for ETO systems.

Reusable launch architecture to support sustainable human exploration of the Solar system

A launch-assisted, horizontal takeoff, reusable Earth-to-orbit (ETO) vehicle is introduced, described, and its concept of operations, a typical trajectory, its performance, mass and cost estimates are presented. This concept enables near aircraft-like operations for a launch system, enhances safety, and reduces cost compared to vertical takeoff systems. Benefits are described. With over a 20 t payload and a payload bay 5 m×15 m , it is compatible with the human exploration architectures presented in a companion paper, IAA.10.3.02. [1]

Highly reusable space transportation: Advanced concepts and the opening of the space frontier

Revolutionary changes in how cargo and people are transported into space are needed to enable the affordable development and exploration of space in the 21st century. Diverse efforts to achieve major, but incremental Earth-to-orbit (ETO) improvements in the relatively near term have been undertaken in recent years in the US, including the Department of Defense evolved expendable launch vehicle system development project. The NASA-industry reusable launch vehicle (RLV) program is addressing this challenge for the mid-term. The RLV program will validate the technology to enable industry to develop all-rocket reusable launch systems that can deliver payloads from the current Civil Needs Data Base in the 20,000–40,000 pounds class and smaller to low Earth orbit (LEO) at costs of approximately $1000–$2000 per pound. This represents a factor of 5 (or more) reduction below existing launch services. This “next generation” improvement in launch capability is a vital element of the US National Space Transportation policy for current and planned government and commercial payloads. The longer-term challenge is also being addressed. During 1995–1997, NASA conducted the highly reusable space transportation (HRST) study project to address the longer-term challenge: how to achieve an additional factor of 10 reduction in launch costs—to approximately $100–$200 per payload pound to LEO—thus enabling a revolutionary expansion of space activity and enterprise. The HRST study has identified a “grand strategy” for achieving these cost goals, based on pursuing a revolutionary advance in main propulsion architectures and technology for ETO systems to enable a dramatic improvements in subsystem operability. The HRST study has examined diverse approaches, including combination propulsion systems, combined cycle propulsion, launch assist systems, and revolutionary rocket propulsion. An integrated assessment has been conducted, including both the concepts defined as part of the study as well as past concepts. This assessment suggests that the cost goals of HRST are achievable within the next 10–20 years if appropriate technology investments are pursued. This paper provides a summary report on the results and findings of the HRST study project.

Health management and controls for Earth-to-orbit propulsion systems

Avionics and health management technologies increase the safety and reliability while decreasing the overall cost for Earth-to-orbit (ETO) propulsion systems. New ETO propulsion systems will depend on highly reliable fault tolerant flight avionics, advanced sensing systems and artificial intelligence aided software to ensure critical control, safety and maintenance requirements are met in a cost effective manner. Propulsion avionics consist of the engine controller, actuators, sensors, software and ground support elements. In addition to control and safety functions, these elements perform system monitoring for health management. Health management is enhanced by advanced sensing systems and algorithms which provide automated fault detection and enable adaptive control and/or maintenance approaches. Aerojet is developing advanced fault tolerant rocket engine controllers which provide very high levels of reliability. Smart sensors and software systems which significantly enhance fault coverage and enable automated operations are also under development. Smart sensing systems, such as flight capable plume spectrometers, have reached maturity in ground-based applications and are suitable for bridging to flight. Software to detect failed sensors has reached similar maturity. This paper will discuss fault detection and isolation for advanced rocket engine controllers as well as examples of advanced sensing systems and software which significantly improve component failure detection for engine system safety and health management.

WARP-10: a numerical simulation model for the cylindrical reconnection launcher

A fully self-consistent computer simulation code called WARP-10, used for modeling the reconnection launcher, is described. WARP-10 has been compared with various experiments with good agreement for performance and heating. Simulations predict that it is possible to obtain nearly uniform acceleration with high efficiency and low armature heating. There does not appear to be an armature heating limit to velocity provided the armature mass can be sufficiently large. Simulation results that show it is possible to obtain conditions needed for earth-to-orbit (ETO) launch applications (4.15 km/s and a 850 kg launch mass) are presented. This 3100-stage launcher has an efficiency of 47.2% and a final ohmic energy/kinetic energy equal to 0.00146. The mode of launcher operation is similar to a traveling wave induction launcher and is produced by properly timed and tuned discrete stages. Further optimization and much higher velocities appear possible. >

Turbine blading designed for high heat load space propulsion applications

The operating conditions and the propellant transport properties used in Earth-to-orbit (ETO) applications affect the aerothermodynamic design of ETO turbomachinery in a number of ways. This paper discusses some aerodynamic and heat-transfer implications of the low molecular weight fluids and high Reynolds number operating conditions on future ETO turbomachinery. The objective of this work was to examine turbine blading concepts for increasing life and reliability by reducing thermal heat load. Using the current Space Shuttle main engine high-pressure fuel turbine as a baseline, aerothermodynamic comparisons were made for two alternate fuel turbine geometries. The first was a revised first-stage rotor blade designed to reduce peak heat transfer. This alternate design resulted in a 23% reduction in peak heat transfer. The second design concept was a single-stage rotor designed to yield the same power output as the baseline two-stage rotor. Since the rotor tip speed was held constant, the turbine work factor doubled. In this alternate design, the peak heat transfer remained the same as the baseline. Although the efficiency of the single-stage design was 3.1 points less than the baseline two-stage turbine, the design was aerothermodynamically feasible and may be structurally desirable.