Rocket Lab Mission to Venus

The dynamics of a small solid launch vehicle has been investigated. This launcher consists of a liquid upper stage and three fundamental solid rocket boosters aligned in series. During the ascent flight phase, lateral jets and grid fins are adopted by the flight control system to stable the attitude of the launcher. The launcher is a slender and aerodynamically unstable vehicle with sloshing tanks. A complete set of six-degrees-of-freedom dynamic models of the launcher, incorporation its rigid body, aerodynamics, gravity, sloshing, mass change, actuator, and elastic body, is developed. Dynamic analysis results of the structural modes and the bifurcation locus are calculated on the basis of the presented models. This complete set of dynamic models can further be used in flight control system design.

Small satellites and the DARPA/Air Force FALCON program

Investigation of different strategies for access to space of small satellites on a defined LEO orbit

Ball Aerospace has recently broadened its line of small satellites as part of the innovative and well-published DARPA `lightsat' initiative. This program, DARPASAT, has produced a small spacecraft bus that is highly capable, lightweight, and can provide significant services to the prospective payload. In addition, the bus is compatible with new small launch vehicles such as the Taurus and Pegasus. Of particular note, the bus-to- payload interface has been designed to maximize simplicity and modularity, all with the intention of easing the payload integration process and thus lowering the cost. As a result, the mechanical, thermal, and electrical interfaces are designed such that the payload can be integrated at the payload contractor's facility as a single entity, or `pallet.' This dramatically reduces the effort to then integrate this single pallet onto the spacecraft bus.

Systemic design of the very-high-resolution imaging payload of an optical remote sensing satellite for launch into the VLEO using an small launch vehicle

Sounding rocket SS-520 as a CubeSat launch vehicle

Elastic-memory composites (EMCs) is a class of polymeric composites that can be folded into a very compact shape and later deployed to the original shape upon reheating. They have significant potential in developing large deployable structures greater than 10m in size that can be stowed in a small launch vehicle and deployed in space to predetermined levels of precision. Generally, the composite is folded by heating it above the glass transition temperature of the resin. The primary deformation mode that allows EMCs to achieve significantly higher effective strains than traditional composites is microbuckling of the compressed fibers. The magnitude of this microbuckling dictates whether the laminate behaves elastically or if the laminate exhibits material failure such as fiber breakage, or fiber kinking. This paper explores the micromechanics of both types of fiber-deformation modes associated with the bending of such soft-resin composites. An analysis method is developed and correlated with experimental data to provide an estimate for the level of folding strain that a soft-resin laminate can achieve based on microbuckling-induced material failure. Additionally, insight is provided into the effects of fiber and matrix material properties on the formation and magnitude of the kink mode. Tests were also conducted to determine the deployment accuracy after a number of folding/deployment cycles to help design deployable sensor structures with these materials.

To meet the ever-expanding needs of satellite customers, a new generation of launch vehicles in the Long-March (LM) family is being developed according to the principles of low-cost, high-reliability, and maximum flexibility. By reducing the pre-launch preparation time, the launch system will strengthen the capability of providing international commercial launch services in the future. The new generation launch vehicles consist of several standard modules featuring two types of newly developed engines and three new modules powered by non-toxic, non-polluting propellants. Among the three modules, 5-m module is powered by two 50-ton-thrust engines which burn liquid hydrogen (LH2) and liquid oxygen (LOX), the 3.35-m module is powered by two 120-ton-thrust engines which burn LOX and kerosene (KO), and the 2.25-m one is powered by one 120-ton-thrust engine. As a first step, a series of large-scale launch vehicles is formed with 5-m-dia common core. By modifying the modules, the next step is to combine medium launch vehicles with the 3.35-m-dia core and a small launch vehicle with the 2.25-m-dia core. Finally a new group of launch vehicles will be formed to accommodate a wide range of performance requirements, which provide a capability of delivering payloads ranging from 1.5 to14 tons to geosynchronous transfer orbit (GTO) and up to 25 tons to low earth orbit (LEO). This paper highlights the design principles, systems concept and a series of configurations for the new generation of the LM launch vehicles. Copyright  2003 by Beijing Institute of Astronautical Systems Engineering. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Released to IAF/IAA/AIAA to publish in all forms. The Long-March family and its history The Long-March (LM) launch vehicle program was initiated in mid 1960s in China. In April 1970, the first launch vehicle LM-1 made its maiden flight and sent DFH-1 satellite into the earth orbit, which indicated that China entered the space era. In the past three decades, the LM family was developed rapidly and formed a series of launch vehicles consisting of LM-2C, LM-2E, LM-3, LM-3A, LM-3B, LM-4, etc. These launch vehicles are suitable for various tasks, such as LEO, SSO and GTO missions. The largest launch capacity of the LM family has reached 9,200 kg for LEO, and 5,200 kg for GTO, which basically meets the demands of various customers. Up to now, the LM family has accomplished 70 launches. From October 1996 to July 2003, 28 consecutive successful flights were made, which established a new record in the LM history. Requirements analysis Although LM family has made eye-catching achievements in many important fields of space transportation technology, the assessment of 1 54th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law 29 September 3 October 2003, Bremen, Germany IAC-03-V.1.04 Copyright © 2003 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Released to IAF/IAA/AIAA to publish in all forms. launch services market clearly shows the need for the LM launch vehicles to be upgraded. − Lift-capability and flexibility need to be increased not only to match the continuing upward trend in terms of GEO communication satellites, but also enable new applications such as constellations, scientific missions and multiphase releases. − Present propellants have to be replaced by non-toxic, non-polluting ones in order to satisfy the demand of the environment protection. − Higher reliability and safety are required for commercial launches, especially for the manned mission. − Launch cost must be significantly reduced for the future market.

Utilizing Small Satellite Launch Vehicle - A Cross-Border Collaboration for Space Mission Australia India Technology Research and Innovation (Space MAITRI) by Space Machines Company and New Space India Limited

In view of the high development cost of the Long March series of launch vehicle and the difficulty of implementing reliability tests, a small sample reliability sampling method based on composite equivalency Bayesian fusion is proposed. First, in order to avoid large amounts of prior data submerging the field data with small sample, a comprehensive use of physical equivalency credibility and data compatibility test are used to fully integrate multi-source test data. Then, according to Bayesian theory, the reliability of the fusion data of the launch vehicle that obeys the normal distribution is statistically verified under the complex assumptions. Finally, considering the experimental cost and the constraints of the two types of risks, a nonlinear constraint programming model for solving the minimum sample size is established.

The use of Solar Electric Propulsion (SEP) is evaluated in the context of a mission to Europa. It is shown that smaller missions making use of SEP can deliver a larger payload on a smaller launch vehicle than is possible with an all -ch emical mission. Large missions (over ~3700 kg) are possible on existing launch vehicles only with SEP systems. The cruise times for SEP missions are generally longer than for all -chemical missions, but depend on launch date. In examining missions to Europa , future planners would be well advised to include SEP in their trade studies. I. Introduction upiter’s moon Europa is one of the most important places in the solar system for future scientific study. Recent evidence 1-4 suggests that Europa may have beneat h its icy surface an ocean of liquid salt water that could potentially support a biosphere. The confirmation of the existence of such a subsurface ocean, and the detection of signs of life on Europa, are two of the most compelling science objectives in pla netary science. Europa has been designated as the most important destination in the outer planets by the NRC Decadal Survey for planetary science 5 . In this work, we examine mission ideas for the future exploration of Europa, and suggest how such a mission (or set of missions) could be done using proven launch vehicles and existing propulsion and other technologies. In this work, we discuss a number of possible missions to Europa. We start with concepts for small missions which would carry a reasonably limit ed payload and could include a lander. Ideas for the design of these missions are discussed in some detail; mission ideas that use solar electric propulsion (SEP) for part of the cruise to Jupiter are examined, as well as missions that use chemical propuls ion only. The point designs for the smaller missions are then scaled to larger mission ideas, with concomitantly higher payloads and capabilities. We will show that for large missions, only payloads delivered in part by a SEP system can be undertaken with a single launch using existing technology.

For over half a century, scientists have contemplated the potential existence of life within the clouds of Venus. Unknown chemistry leaves open the possibility that certain regions of the Venusian atmosphere are habitable. In situ atmospheric measurements with a suite of modern instruments can determine whether the cloud decks possess the characteristics needed to support life as we know it. The key habitability factors are cloud particle droplet acidity and cloud-layer water content. We envision an instrument suite to measure not only the acidity and water content of the droplets (and their variability) but additionally to confirm the presence of metals and other non-volatile elements required for life’s metabolism, verify the existence of organic material, and search for biosignature gases as signs of life. We present an astrobiology-focused mission, science goals, and instruments that can be used on both a large atmospheric probe with a parachute lasting about one hour in the cloud layers (40 to 60 km) or a fixed-altitude balloon operating at about 52 km above the surface. The latter relies on four deployable mini probes to measure habitability conditions in the lower cloud region. The mission doubles as a preparation for sample return by determining whether a subset of cloud particles is non-liquid as well as characterizing the heterogeneity of the cloud particles, thereby informing sample collection and storage methods for a return journey to Earth.

A Flagship-class Uranus Orbiter and Probe mission concept using aerocapture

We construct a simple cloud model for a Venus general circulation model (GCM), which includes condensable gases of H2O and H2SO4 vapors, and condensation, evaporation, and sedimentation of sulfuric acid cloud particles. The zonally averaged mass loading of the cloud reproduced in the model takes its maximum and minimum in high and middle latitudes, respectively. This latitudinal distribution is consistent with the infrared measurements. The thick cloud is formed in high latitudes at 43–55 km altitudes by vertical winds associated with disturbances enhanced in the low static stability layer. The moderately thick cloud in low latitudes is attributed mainly to the transport of H2SO4 vapor by the mean meridional circulation. The horizontal cloud distribution in low latitudes has zonal wave numbers 1 and 2 structures, which change in time significantly. These characteristics of the low‐latitude cloud would be associated with atmospheric waves in the cloud layer. The mixing ratio of H2O vapor increases with latitude in the cloud layer due to the vertical wind disturbances in the low static stability layer in high latitudes. This latitudinal trend is qualitatively consistent with the infrared measurements. The mixing ratio of H2SO4 vapor increases with latitude in the subcloud layer because a large amount of the cloud is evaporated there due to the sedimentation of cloud particles in the thick lower cloud in the polar region. The present results suggest that the Venus cloud distribution in the lower cloud layer is strongly affected by waves and/or disturbances as well as the mean meridional circulation.

A research program has been initiated at NASA Langley Research Center to investigate the critical technologies for developing advanced microwave radiometers suitable for Earth science observations. A significant objective of this research is to enable microwave measurements with adequate spatial resolutions for a number of Earth science parameters, such as sea ice, precipitation, soil moisture, sea surface temperature, and wind speed over oceans. High spatial resolution microwave sensing from space with reasonable swath widths and revisit times favor large real aperture radiometer systems. However, the size requirements for such systems are in conflict with the need to emphasize small launch vehicles. This paper describes a tradeoff between the science requirements, basic operational parameters, test configurations, and expected sensor performance for a satellite radiometer concept. The preliminary designs of real aperture systems utilizing novel light-weight compact-packaging techniques are used as a means of demonstrating this technology.