Future Launch Vehicles Research Articles

Hybrid boosters for future launch vehicles

There is a striking similarity in the design of the US Space Transportation System, the European ARI-ANE 5P and the Japanese II-II: they all use a high energy cryogenic core stage along with two large solid propellant rocket boosters (SRB's) in order to provide for a high lift-off thrust level. Prior to last years disasters with Challenger and Titan it was widely held that SRB's were cheap, uncomplicated and safe. Even before the revelation by these accidents of severe safety hazards, shuttle operations demonstrated that the SRB's were by no means as cheap as reusable systems ought to be. In addition, they became known as sources of considerable environmental pollution. In contrast, hybrid rocket propulsion systems offer the following potential advantages: • • much higher savety level (TNT equivalent almost zero, shut-down capability in case of ignition failure of one unit, inert against unbonding) • • choice of non-toxic propellant combinations • • equal or higher specific performance For these reasons, system analysis were carried out to examine hybrids as potential alternative to SRB's. Various analytical tools (mass- and performance models, trajectory simulation etc.) were developed for parametrical studies of hybrid propulsion systems. Special attention was devoted to the well-known primary concern of hybrids: geometrical design of the solid fuel grain and regression rate of the ablating surface. Experimental data were used as input wherever possible. In 1985 first studies were carried out to find possible fields of application for hybrid rocket engines. A mass model and a performance model for hybrid rocket motors were developed, taking into account the peculiarities of hybrid combustion as there are i.e. low regression rate and shifting mixture ratio during operation. After some analytical work was done, hybrids proved to be a promising alternative to SRB's. Compared with solids, hybrids offer many advantages.

Assessment of existing and future launch vehicle liquid engine development

Existing liquid propellant engines for large launch vehicles are described in terms of pertinent engine and propellant parameters and their launch vehicle application. The development approach and the maturity of engine technology which prevailed prior to and early in specific engine development programs are discussed including lessons learned. New engines, including improved conventional and new concepts that could support the next generation launch vehicle, are delineated with emphasis on technology. The technology maturity and development needed to alleviate the potential development risk are presented along with projected gains in performance, operability, reusability, reliability and producibility. A technology ranking methodology which incorporates a relative transportation system life cycle cost (LCC) as the payoff function is developed. The methodology is useful for establishing preliminary but timely cost effectiveness rankings of various technologies. The methodology uses conventional cost estimating relations (CER) in non-dimensional form. The relative overall transportation system payoff resulting from the implementation of new propulsion system technology is developed from recurring and non-recurring costs of major transportation system elements including the vehicle, operations and launch complex. A ranking of engine concepts and associated technologies is given for several transportation system candidates which serve a high activity mission model.

Advanced propulsion activities in the USA

Advances in propulsion system performance are necessary to meet the projected requirements of future launch vehicles, orbit transfer vehicles, satellites and planetary spacecraft. Ongoing U.S. space propulsion research programs are aggressively pursuing the development of the technology to meet these requirements. This paper reviews advanced propulsion research in the USA. Work accomplished in 1985 and future plans for ongoing research programs are discussed. Theoretical and experimental research programs and system studies of augmented hydrazine, arcjet, magnetoplasma dynamic thruster, ion thruster, metastable chemical, solar thermal, light sail, microwave and laser propulsion, nuclear fission rockets, and fusion and antimatter propulsion systems are addressed.

Development tests of LOX/LH 2 tank for H-I launch vehicle

H-I is a future launch vehicle of Japan with a capability of placing more than 550 kg payload into a geostationary orbit. The National Space Development Agency of Japan (NASDA) is now directing its efforts to the final development of H-I launch vehicle. H-I's high launch capability is attained by adopting a newly developed second stage with a LOX/LH 2 propulsion system. The second stage propulsion system consists of a tank and an engine. The tank is 2.5 m in diameter and 5.7 m in length and contains 8.7 tons of propellants. This tank is an integral tank with a common bulkhead which separates the tank into forward LH 2 tank and aft LOX tank. The tank is made of 2219 aluminum alloy and is insulated with sprayed polyurethane foam. The common bulkhead is made of FRP honeycomb core and aluminium alloy surface sheets. The most critical item in the development of the tank is the common bulkhead, therefore the cryogenic structural test was carried out to verify the structural integrity of the bulkhead. The structural integrity of the whole LOX/LH 2 tank was verified by the cryogenic structural test of a sub-scale tank and the room temperature structural test of a prototype tank.

Liftoff Ignition Overpressure-A Correlation

Current and planned launch vehicle configurations consist of extremely large vehicles requiring high-thrust propulsion systems for boost. A critical design condition for this type of vehicle is the pressure environment caused by the ignition start transient. The vehicle and surrounding support facilities must be designed to withstand these conditions. The resulting pressure wave emanating from the launch duct entrance can travel the entire length of the launch vehicle. This paper shows the results of full-scale measurements on several Titan vehicle launch conditions. Subscale model (Titan III) tests were also evaluated at a 7.5% model scale by NASA. The resultant correlation of subscale with full-scale data produces a scaling parameter that should be useful to vehicles such as shuttle and future launch vehicles with even larger thrust potentials.

The impact of launch vehicle type and size on development cost

The technical development trend of future launch vehicle systems is towards fully reusable systems, in order to reduce space transportation cost. However, different types of launch vehicles are feasible, as there are • —winged two-stage systems (WTS) • —ballistic single-stage vehicles (BSS) • —ballistic two-stage vehicles (BTS) The performance of those systems is compared according to the present state of the art as well as the development cost, based on the “TRANSCOST-Model”. The development costs are shown versus launch mass (GLOW) and pay-load for the three types of reusable systems mentioned above. It is shown that performance optimization and cost minimization lead to different results. It is more economic to increase the vehicle size for achieving higher performance, instead of increasing technical complexity. Finally it is described that due to the essentially lower launch cost of reusable vehicles it will be feasible to recover the development cost by an amortization charge on the launch cost. This possibility, however, would allow commercial funding of future launch vehicle developments.

Economic considerations of the determination of investments for launch vehicle costs in the next two decades (1980–2000)

Investments in launch vehicle costs will be substantial in the next two decades, both in the RDT & E as well as the investment/operation phases. For Europe the “investment”; question is further complicated by the choice of (a) making full use of the Space Shuttle system or (b) further expansion and development of Shuttle independent space transportation capabilities. Considerations on future launch vehicle costs will be increasingly resolved on purely economic grounds: are the proposed investments economically efficient in the sense of the best use of these funds for space activities? Investments for reasosn of “prestige”; or “technology push” are probably better spent on payload programs, rather than space transportation systems. Economic considerations played a major factor in the choice of the space Shuttle system being developed in the United States. The economic considerations applied were: • • A comparison of space transportation systems (STS) costs of expendable vs reusable STSs of about $21 billion each (May 1979 dollars) • • The effects of different STSs on payload costs, including (1) RDT & E costs, (2) investment and operations costs and (3) on assured reliability of programs. (Repair, refurbishment, check-up capabilities) of about $63 billion (STS) and $76 billion (Expendables) respectively. These effects of STS on payload costs ($13 billion) often are lumped together under the term “payload effects”. They proved decisive in the choice of a Space Shuttle system, and the precise configuration now being developed and completed by NASA. Three economic evaluation approaches can be used separately or jointly when comparing space transportation system investments: • • Equal capability analyses • • Equal budget analyses • • New capability analyses. Looking ahead to investments into further launch vehicle developments in the 1980s and 1990s the same criteria will apply to any economic choice among rather expensive programs.

Development of launch vehicles for application purposes

The N-I, N-II, H-I launch vehicles and TT-500A rocket will be described in this paper out of the various launch vehicles and rockets developed and being developed by the National Space Development Agency of Japan (NASDA). The N-I is at present the main satellite launcher of Japan. The N-II is an improved version of N-I, and the H-I is a future launch vehicle with a cryogenic second stage. The TT-500A is a small two stage rocket for material processing experiments. Since the first stage of NASAs Thor-Delta vehicle is used as the first stage of N-I and N-II, most of the development efforts have been expended on the second stages. The N-I, N-II and H-I second stages are presented to show the development activities at NASDA.

Future trends in communications satellite systems

This paper presents a survey of the progress and future trends in communications satellite technology which are expected to contribute most significantly to the rapidly expanding use of this medium for international and domestic communications including fixed and mobile services. Future launch vehicle capabilities and spacecraft, earth terminal, and communications processing technologies are considered. Systems implications of trends in these areas are related to cost effectiveness, orbital utilization, and spectrum sharing.

Launch vehicles for future automated space missions.

Future launch vehicle programs for automated space missions, discussing higher launch velocities and economics

Evaluation of multimission launch vehicle concepts.

The use of future launch vehicles for a multitude of mission applications appears to he a necessity. The specific applications for these future systems are uncertain at this time. The launch rates for present and projected future missions are relatively low. These low launch rates and the uncertainties in performance and operational modes focus attention on launch vehicles that have the potential of performing a range of missions. An evaluation and comparison is sho\vn for multimission and single-mission concepts capable of performing three typical future missions. This paper uses 13 evaluation criteria to measure the effectiveness of the several launch concepts. The results of a consensus to establish the relative value of the criteria are discussed. The flexibility of the multimission vehicles is shown to be achievable with only a small penalty in weight, although a superiority in cost is maintained.