Liquid rocket engine test facility engineering challenges

Liquid rocket engines for launch vehicles and space crafts as well as their subsystems need to be verified and qualified during hot-runs. A high test cadence combined with a flexible test team helps to reduce the cost for test verification during development/qualification as well as during acceptance testing for production. Test facility intelligence allows to test subsystems in the same manner as during complete engine system tests and will therefore reduce development time and cost.This paper gives an overview of the maturing of test engineering know how for rocket engine test stands as well as high altitude test stands for small propulsion thrusters at EADS-ST in Ottobrunn and Lampoldshausen and is split into two parts:

• Part 1 gives a historical overview of the EADS-ST test stands at Ottobrunn and Lampoldshausen since the beginning of Rocket propulsion activities in the 1960s.

• Part 2 gives an overview of the actual test capabilities and the test engineering know-how for test stand construction/adaptation and their use during running programs.

Examples of actual realised facility concepts are given to demonstrate cost saving potential for test programs in both cases for development/qualification issues as well as for production purposes.

Piloted operations at a near-Earth object (NEO)

In late 2006, NASA's Constellation Program sponsored a study to examine the feasibility of sending a piloted Orion spacecraft to a near-Earth object. NEOs are asteroids or comets that have perihelion distances less than or equal to 1.3 astronomical units, and can have orbits that cross that of the Earth. Therefore, the most suitable targets for the Orion Crew Exploration Vehicle (CEV) are those NEOs in heliocentric orbits similar to Earth's (i.e. low inclination and low eccentricity). One of the significant advantages of this type of mission is that it strengthens and validates the foundational infrastructure of the United States Space Exploration Policy and is highly complementary to NASA's planned lunar sortie and outpost missions circa 2020. A human expedition to a NEO would not only underline the broad utility of the Orion CEV and Ares launch systems, but would also be the first human expedition to an interplanetary body beyond the Earth–Moon system. These deep space operations will present unique challenges not present in lunar missions for the onboard crew, spacecraft systems, and mission control team. Executing several piloted NEO missions will enable NASA to gain crucial deep space operational experience, which will be necessary prerequisites for the eventual human missions to Mars.Our NEO team will present and discuss the following:

• new mission trajectories and concepts;

• operational command and control considerations;

• expected science, operational, resource utilization, and impact mitigation returns; and

• continued exploration momentum and future Mars exploration benefits.

Microgravity research and user support in the Space Station era: The Microgravity User Support Centre

Future space systems, such as Columbus, the planned European contribution to the International Space Station, offer ample possibilities for microgravity research and application. These new opportunities require adequate user support on ground and novel operational concepts in order to ensure an effective utilization. Extensive experience in microgravity user support has been accumulated at DFVLR during the past Spacelab 1 and D1 missions. Based on this work, a Microgravity User Support Centre (MUSC) has been built and is active for the forthcoming EURECA-A1 and D2 missions, to form an integrated support centre for the disciplines life sciences and material sciences in the Space Station era. The objective of the user support at MUSC is to achieve:

• easy access to space experiments for scientific and commercial users,

• efficient preparation of experiments,

• optimum use of valuable microgravity experimentation time,

• cost reduction by concentration of experience.

This is implemented by embedding the MUSC in an active scientific environment in both disciplines, such that users can share the experience gained by professional personnel. In this way, the Space Station system is operated along the lines established on ground for the utilization of large international research facilities, such as accelerators or astronomical observatories. In addition, concepts are developed to apply advanced telescience principles for Space Station operations.     

Future launchers strategy : the ariane 2010 initiative

With the new cryogenic upper stage ESC, the European heavy launcher Ariane 5+ is perfectly suited to the space market envisioned for the coming decade: flexible to cope with any payload and commercially attractive despite a fierce competition.Current Arianespace projections for the following years 2010–2020 indicate two major trends:

• satellites may still become larger and may require very different final orbits; today's market largely dominated by GEO may well evolve, influenced by LEO operations such as those linked to ISS or by constellations,

• to remain competitive, the launch cost has to be reduced.

The future generation of the European heavy launcher has therefore to focus on an ever increased flexibility with a drastic cost reduction.Two strategies are possible to achieve this double goal:

• reusable launchers, either partially or totally, may ease the access to space, limiting costly expendable stages; the assessment of their technical feasibility and financial viability is undergoing in Europe under the Future Launchers Technology Program (FLTP),

• expendable launchers, derived from the future Ariane 5+.

This second way started by CNES at the end of year 1999 is called the “Ariane 2010 initiative”.The main objectives are simultaneously an increase of 25% in performance and a reduction of 30% in launch cost wrt Ariane 5+.To achieve these very ambitious goals, numerous major modifications are studied:

• technical improvements :

◦ modifications of the Solid Rocket Boosters may consist in filament winding casing, increased loading, simplified casting, improved grain, simplified Thrust Vector Control, …

◦ evolution of the Vulcain engine leading to higher efficiency despite a simplified design, flow separation controlled nozzle extension, propellant management of the two cryogenic stages,

◦ simplified electrical system,

◦ increased standardization, for instance on flanged interfaces and manufacturing processes,

• operational improvements such as launch cycle simplification and standardization of the coupled analyses,

• organizational improvements such as a redistribution of responsibilities for the developments.

All these modifications will of course not be implemented together; the aim is to have a coherent catalogue of improvements in order to enable future choices depending on effective requirements. These basic elements will also be considered for the development of other launchers, in the small or medium size range.     

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.     

Space solar arrays and concentrators

This paper presents some research activities conducted at the Centre Spatial de Liege (CSL) in the field of space solar arrays and concentration.With the new generation of high efficiency solar cells, solar concentration brings new insights for future high power spacecrafts. A trade-off study is presented in this paper. Two different trough concentrators, and a linear Fresnel lens concentrator are compared to rigid arrays. Thermal and optical behaviors are included in the analysis.Several technical aspects are discussed:

• Off-pointing with concentrators induces collection loss and illumination non uniformity, reducing the PV efficiency.

• Concentrator deployment increases the mission risk.

• Reflective trough concentrators are attractive and already proven. Coating is made of VDA (Aluminum). A comprehensive analysis of PV conversion increase with protected silver is presented.

• Solar concentration increases the heat load on solar cells, while the conversion efficiency is significantly decreasing at warm temperatures.

To conclude, this paper will point out the new trends and the key factors to be addressed for the next generation of solar generators.     

Advanced technologies available for future solid propellant grains

Significant advances have been made during the last decade in several fields of solid propulsion: the advances have enabled new savings in the motor development phase and recurring costs, because they help limit the number of prototypes and tests.The purpose of the paper is to describe the improvements achieved by SNPE in solid grain technologies, making these technologies available for new developments in more efficient and reliable future SRMs: new energetic molecules, new solid propellants, new processes for grain manufacturing, quick response grain design tools associated with advanced models for grain performance predictions.Using its expertise in chemical synthesis, SNPE develops new molecules to fit new energetic material requirements.Tests based on new propellant formulations have produced good results in the propellant performance/safety behavior ratio. New processes have been developed simultaneously to reduce the manufacturing costs of the new propellants.In addition, the grain design has been optimized by using the latest generation of predictive theoretical tools supported by a large data bank of experimental parameters resulting from over 30 years' experience in solid propulsion:

• Computer-aided method for the preliminary grain design

• Advanced models for SRM operating and performance predictions

Planetary Defense From Space: Part 2 (Simple) Asteroid Deflection Law

A system of two space bases housing missiles for an efficient Planetary Defense of the Earth from asteroids and comets was firstly proposed by this author in 2002. It was then shown that the five Lagrangian points of the Earth–Moon system lead naturally to only two unmistakable locations of these two space bases within the sphere of influence of the Earth. These locations are the two Lagrangian points L1 (in between the Earth and the Moon) and L3 (in the direction opposite to the Moon from the Earth). In fact, placing missiles based at L1 and L3 would enable the missiles to deflect the trajectory of incoming asteroids by hitting them orthogonally to their impact trajectory toward the Earth, thus maximizing the deflection at best. It was also shown that confocal conics are the only class of missile trajectories fulfilling this “best orthogonal deflection” requirement.The mathematical theory developed by the author in the years 2002–2004 was just the beginning of a more expanded research program about the Planetary Defense. In fact, while those papers developed the formal Keplerian theory of the Optimal Planetary Defense achievable from the Earth–Moon Lagrangian points L1 and L3, this paper is devoted to the proof of a simple “(small) asteroid deflection law” relating directly the following variables to each other:

(1) the speed of the arriving asteroid with respect to the Earth (known from the astrometric observations);

(2) the asteroid's size and density (also supposed to be known from astronomical observations of various types);

(3) the “security radius” of the Earth, that is, the minimal sphere around the Earth outside which we must force the asteroid to fly if we want to be safe on Earth. Typically, we assume the security radius to equal about 10,000 km from the Earth center, but this number might be changed by more refined analyses, especially in the case of “rubble pile” asteroids;

(4) the distance from the Earth of the two Lagrangian points L1 and L3 where the defense missiles are to be housed;

(5) the deflecting missile's data, namely its mass and especially its “extra-boost”, that is, the extra-energy by which the missile must hit the asteroid to achieve the requested minimal deflection outside the security radius around the Earth.

This discovery of the simple “asteroid deflection law” presented in this paper was possible because:

(1) In the vicinity of the Earth, the hyperbola of the arriving asteroid is nearly the same as its own asymptote, namely, the asteroid's hyperbola is very much like a straight line. We call this approximation the line/circle approximation. Although “rough” compared to the ordinary Keplerian theory, this approximation simplifies the mathematical problem to such an extent that two simple, final equations can be derived.

(2) The confocal missile trajectory, orthogonal to this straight line, ceases then to be an ellipse to become just a circle centered at the Earth. This fact also simplifies things greatly. Our results are thus to be regarded as a good engineering approximation, valid for a preliminary astronautical design of the missiles and bases at L1 and L3.

Still, many more sophisticated refinements would be needed for a complete Planetary Defense System:

(1) taking into account many perturbation forces of all kinds acting on both the asteroids and missiles shot from L1 and L3;

(2) adding more (non-optimal) trajectories of missiles shot from either the Lagrangian points L4 and L5 of the Earth–Moon system or from the surface of the Moon itself;

(3) encompassing the full range of missiles currently available to the USA (and possibly other countries) so as to really see “which missiles could divert which asteroids”, even just within the very simplified scheme proposed in this paper.

In summary: outlined for the first time in February 2002, our Confocal Planetary Defense concept is a simplified Keplerian Theory that already proved simple enough to catch the attention of scholars, popular writers, and representatives of the US Military. These developments would hopefully mark the beginning of a general mathematical vision for building an efficient Planetary Defense System in space and in the vicinity of the Earth, although not on the surface of the Earth itself!We must make a real progress beyond academic papers, Hollywood movies and secret military plans, before asteroids like 99942 Apophis get close enough to destroy us in 2029 or a little later.     

Potential cost reductions for three STS/Spacelab mission modes

Based on the results of studies carried out by ESA several possibilities are discussed to achieve mission cost reductions for large Spacelab instrument facilities as compared to their flight on several 7-day duration Spacelab missions. As an example three scientific telescope facilities are selected (LIRTS, EXSPOS, GRIST) which are defined to a Phase A level.Three new mission modes are considered:

• —Shuttle attached Spacelab mission mode with extended flight duration (up to 30 days) for which the application of planned capability extensions and new elements of the STS/Spacelab (e.g. Short Spacelab Pallets, Power Extension Package) are investigated.

• —Shuttle deployed mission mode, for which the telescope, accommodated on a Spacelab pallet, is docked to the Power Module, a new element of the Space Transportation System under study by NASA.

• —Free-flying mission mode, for which Shuttle launched dedicated missions of the facilities are considered, assuming varying degrees of autonomy with respect to supporting services of the Shuttle.

Reduction of costs have been considered on the levels of single mission cost and total programme cost. Fundamentally the charges for the instrument can be reduced by constraining the mass/volume factors with respect to the Shuttle capability. However, the instrument as part of a payload is only viable if an acceptable resource sharing including observation time can be achieved. Any single instrument will require several mission opportunities or one mission which achieves a similar or longer total observation programme.Based on an identification of instrument modifications of the Phase A baseline designs to favour cost reductions and on a derivation of technical requirements, constraints and finally budgetary cost comparisons an attempt is made to assess the advantages and disadvantages of the different mission modes.The favoured option for GRIST is a 2–3 weeks sortie mission followed after refurbishment by a longer Power Module docked mission. For LIRTS and EXSPOS the free-flying pallet modes are very attractive in terms of the longer durations achieved and in terms of cost per unit operating time.