Transcribed from the European Space Agency's 

 Quarterly Publication on Space Transportation Systems

         Volume 2, Number 6, dated July 1991


 ****> From the Director's Chair

 Space transportation remains a focus of European space
 activities. Ariane-4 is back on the road of success: after
 V36, seven successful launches have been performed. Ariane-5
 is well on its track for the qualification launches planned in
 1995. The first full thrust tests with a complete Vulcain
 engine have been successful, the completion of the launch and
 production facilities in Kourou is progressing very
 satisfactorily. The first batch of solid propellant has
 recently been mixed in the new solid propellant facility. In
 Les Mureaux, the launcher integration facilities have recently
 been inaugurated.

 Using the results of the work performed by industry and
 research organisations in the framework of phase 1 of the
 Hermes development programme, the Agency is presently
 completing its proposals for the Council at ministerial level,
 which will take place in Germany in November 1991. We are
 confident that a positive decision for phase 2 of this
 important development programme will be achieved, allowing all
 of us to make a big step forward in the field of European
 manned space transportation and to demonstrate the
 technological ability and the political determination of
 Europe in this area. Due to financial restrictions, the
 schedule of the Hermes development programme has been
 rearranged to lead to a first launch in the year 2000,
 followed by a second qualification launch in 2001. The
 development programme will be followed by an operational
 validation programme that includes two additional missions.
 The exploitation programme will then lead to a first flight
 servicing of the Columbus Free-Flyer in 2004. Even though this
 revised planning appears at a first glance to be rather
 stretched out and demanding, it requires, nevertheless, from
 all of us, a considerable effort and discipline to master all
 aspects of this challenging programme.

 J. Feustel-Buechl
 Director
 Space Transportation Systems


 ***> ESA Studies on Future Launchers

 **> The Near-Term Reusable Rocket Launcher

 ===> In the previous issue of this Newsletter, the scope of
 the ESA studies on future launchers was outlined. In this
 continuation, we begin our presentation of the solutions under
 investigation with the Near-Term Reusable Rocket Launcher, RRL
 for short. By our definition, the RRL is a two-stage to low
 Earth orbit (LEO) rocket like Ariane-5, but with liquid rocket
 propulsion embodied in both stages and with a recoverable
 first stage. Reusable launchers are studied with the hope that
 specific launch costs (the cost of launching one kilogramme of
 payload into orbit) below that of a fully expendable launcher
 (where all stages are discarded after their propulsion phase)
 could be achieved.

 Making a stage reusable is a technical complication, and it is
 not yet clear under which conditions the savings expected from
 multiple reuses could outweigh the additional development
 costs. A first round of studies on this theme was performed in
 the period 1980 to 1984, with Aerospatiale as prime
 contractor, under the generic heading of FLS (Future Launching
 Systems) studies.

 At that time, the Ariane-5 launcher was still in its
 definition phase, and its architecture and propellant choices
 were still open. These FLS studies concluded that the optimal
 number of stages to reach LEO was two, that the reuse of the
 1st stage was certainly feasible with near-term technology and
 entailed only a small performance penalty, that the reusable
 1st stage could be powered by rockets burning liquid oxygen
 with methane or with liquid hydrogen, and that the 2nd stage
 must be a liquid oxygen/liquid hydrogen stage, the reuse of
 which would be very penalising unless advanced technologies
 could be employed. The FLS study concluded with an optimal
 architecture which puts the two stages in series, and in which
 only the first stage is reusable.

 This so-called semi-reusable FLS was shown to be of economical
 interest when compared with an entirely expendable launcher of
 same propulsive technology. However, by 1984, the present
 design of Ariane-5 was decided (parallel architecture and
 solid propellant boosters) and the FLS was found to offer only
 about a 15% lower recurrent cost than the Ariane-5
 configuration. With the decision to develop Ariane-5, the
 studies on the FLS were then shelved in late 1984.

 It is interesting to note that, shortly after that time, a
 significant study effort was begun in the USA under the term
 of STAS (Space Transportation Architectures Studies) which
 eventually led to the ALS concept (the Advanced Launching
 System), which itself is now evolving into the ALDP (Advanced
 Launcher Development Programme), aiming at a heavy lift launch
 capability at affordable cost and with high reliability. It is
 of interest to note that the launcher architecture identified
 as the most economical one in the context of the STAS was
 similar to the semireusable FLS described above.


 Today, the development of Ariane-5 is in full progress and ESA
 has again taken up the subject of reusable rocket launchers by
 linking it very closely to the technology base that will be
 provided by the Ariane-5 programme. Of particular relevance
 are the cryotechnic propulsion elements that are being
 developed (the Vulcain engine, the cryogenic storage tank -
 H150) and the facilities being constructed for the assembly
 and launch of Ariane-5 (the ELA-3). The recent studies that
 have been done on this theme will be presented in a
 forthcoming article.

 H. Pfeffer
 ESA

 ***> MBB's participation in the development of the cryogenic
 stage

 ===> MBB plays an important role in the development and
 fabrication of the Vulcain motor. The development and
 production of the thrust chamber, the heart of the Vulcain
 motor, as well as the development testing of the Liquid Oxygen
 turbopump (subcontracted by Fiatavio) and of the gas generator
 (subcontracted by SEP) are under MBB responsibility.
 Furthermore, numerous valves of the cryogenic main stage are
 subcontracted to MBB, such as the main feed valves, the tank
 pressurisation units, the tank valves and the valves of the
 anti-pogo device and of the liquid helium feed.

 The thrust chamber firings started in December 1988 on the
 P3.2 test facility in Lampoldshausen that was designed, built
 and commissioned by MBB from 1984 to 1989. Since then, over 90
 successful hot and cold tests were performed on this test
 facility where the motor is mounted with its axis horizontal.
 The maximum run duration is limited by the tank capacity to 20
 seconds. The thrust chamber has already demonstrated its
 capabilities in the entire operational and extreme envelope.
 Since 1989 MBB has delivered five thrust chambers to SEP for
 motor integration and testing. Motor testing has started in
 mid-1990 and since then the thrust chamber has undergone about
 35 tests at engine level. The gas generator tests and the LOx
 turbopump tests are performed on MBB's P59 test facility in
 Ottobrunn. Up to now some 120 successful tests have been
 performed, providing important data for design improvements
 and information on the operation of these motor components.

 E. Kirner
 DASA-MBB


 ***> Successful payload fairing cylinder separation test
 performed at Oerlikon-Contraves

 ===> The satisfactory performance of the payload fairing
 separation systems was demonstrated on 22 March with the first
 functional separation test performed at Oerlikon-Contraves,
 the company responsible for the payload fairing development.
 This was an important development step providing confidence in
 the design before the start of the qualification programme.

 The lower cylindrical part of the payload fairing, with a
 diameter of 5.4 m and a height of 2 m, was used in the test.
 The test was performed in normal atmospheric conditions.

 Separation and jettison from the launcher is achieved by means
 of two pyrotechnically operated separation systems which had
 never previously been tested together, nor tested with the
 full-size cylindrical carbon-fibre sandwich structure of 5.4 m
 diameter.

 One of the two systems, the horizontal separation system
 developed by Aerospatiale, separated the fairing from the
 launcher. Separation into two fairing halves and lateral
 jettison from the launcher is carried out by the Oerlikon-
 Contraves vertical separation system.

 Full separation is achieved within a few milliseconds. The
 critical clearance distance between the fairing and the
 launcher is reached at 0.5 s after jettison and is, to a large
 extent, determined by the breathing, opening and closing, of
 the half shells. A second functional cylinder separation test
 will take place this year following evaluation of the test
 results. The full-scale separation tests in a vacuum chamber
 with a 12.7 m long payload fairing are foreseen for the end of
 1992.

 For extra large payloads, such as the Columbus elements, a
 second version with a length of 19.2 m is also being
 developed. The experience of Oerlikon-Contraves in developing
 payload fairings, and the separation test results obtained
 with the 12.7 m long version will enable the long payload
 fairing to be qualified using analysis predictions only
 despite its different separation behaviour.

 An essential increase in the payload mass capability of
 Ariane-5 was achieved with the aerodynamically favourable,
 although technological challenging, ogive front part which was
 selected based on aerodynamic studies and wind tunnel tests.
 The payload mass capability for a geostationary transfer orbit
 was thereby increased by 100 kg and for a low Earth orbit by
 200 kg.

 Manufacturing of the first payload fairing qualification test
 unit has started. Final integration of this test unit will be
 completed in the beginning of 1992 in order to be ready for
 the qualification programme which will start with the
 performance of the static load tests.


 J.P. Schwander
 Oerlikon-Contraves AG


 ***> ARIANE-5 and the Ozone Layer: An expert's View

 ===> There is an increasing worldwide concern for the
 protection of our natural environment. This has included
 suggestions from various persons in Europe, in the Soviet
 Union and even in the United States, that the US Space Shuttle
 and Ariane-5 could cause damage to the environment. This
 situation prompted the European Space Agency to call on the
 latest results of atmospheric research to determine whether
 the concern was indeed justified, and to determine in which
 directions additional investigations might be desirable.

 The study was focused on the implementation of large solid
 propellant boosters burning a propellant based on ammonium
 perchlorate, since this is a novel feature of the Ariane-5
 programme. Before reporting on the results of this work it is
 necessary to recall briefly what the ozone problem is.

 **> The Ozone Problem

 Ozone is a gas found in the atmosphere in very small
 quantities; it is distributed all through the atmosphere but
 its highest relative abundance is in the stratosphere, at an
 altitude between 20 and 35 km. A decrease in this abundance
 would lead to a decreased effectiveness of the ozone shield
 against UV radiation. Changes in ozone can also lead to
 changes in the atmospheric temperature profile; changes in
 ozone in the lower stratosphere could have a particularly
 important impact on surface temperature. Thus ozone change is
 considered to be potentially very significant, due to its
 impact on weather patterns and on climate.

 Ozone is created and destroyed in the upper atmosphere:
 creation is by reaction activated by UV radiation, destruction
 is by various reactions involving, in particular, the oxides
 of hydrogen, nitrogen, chlorine and bromine. The result is an
 equilibrium. The natural chlorine that plays a role is
 released by the oceans at the rate of about 300 million tonnes
 per year. Other chlorine compounds are also naturally produced
 at the surface, sometimes in large amounts but, because of
 their high solubility, they are rapidly washed out in the
 lower atmosphere. They do not reach the stratosphere to
 contribute to ozone destruction. For example, usually only a
 negligible fraction of the 60 million tonnes of chlorine
 produced per year by volcanoes reaches the stratosphere.

 The delicate balance has been upset by man by the increased
 emissions of chlorofluorocarbons (CFC's). These substances are
 characterised by an extraordinary stability against
 destruction in the troposphere, against wash out by water and
 decomposition by 'soft' UV radiation. They have therefore a
 long lifetime in the atmosphere (up to 10 - 100 years) and
 tend to accumulate. They are carried into the upper atmosphere
 and into the ozone layer where they find harder UV radiation
 that decomposes them, liberating the destructive chlorine. At
 the core of the ozone problem is the fact that CFC's which are
 released at the rate of about 800,000 tonnes per year, release
 their chlorine in the stratosphere.

 **> Ariane-generated Gases

 Solid propellant rocket motors generate hydrochloric acid that
 is ejected from the nozzle as the vehicle moves through the
 atmosphere. The quantities involved are 0.2 tonne of
 hydrochloric acid for 1 tonne of solid propellant. For
 Ariane-5 launches at a rate of 10 per year, this means the
 generation of 960 tonnes of hydrochloric acid that can break
 down into chlorine. This chlorine is chemically similar to
 that generated by the seas and the volcanoes (it is
 not as inert as the chlorine bound in the CFC molecules) and
 the contribution of the rocket flights must be compared to the
 approximately 300 million tonnes of chlorine of natural
 origin.

 The contribution of the rocket motors is so small that effects
 in the troposphere can be discounted, only local effects must
 be considered and are actually taken into account in the
 management of the launch site.

 In reality the solid propellant burns partly in the lower
 atmosphere and partly in the upper atmosphere. From the
 Ariane-5 trajectory one can determine that about half of the
 solid propellant is burned in the lower part of the
 atmosphere, the troposphere (up to 20 km altitude), the other
 half being burned and exhausted in the stratosphere, where the
 chlorine released could affect ozone. Since no experimental
 nor theoretical results on the effect of 'in-situ' release of
 chlorine in the stratosphere were known to ESA, it was decided
 to look in more detail at this particular problem. The problem
 has been analysed by Dr J. Pyle of the European Ozone
 Secretariat, at the University of Cambridge, and the British
 Antarctic Survey, where the ozone hole was discovered. He has
 shown that if Ariane-5 launches are performed from Kourou at
 the rate of 10 per year over a period of 20 years, the
 chlorine content of the ozone layer would increase locally by
 approximately 0.35%, and the ozone content would decrease
 locally by 0.1%. These are steady state values, where a new
 balance between production and destruction of ozone is
 achieved. These local values are very small. Values averaged
 over the stratosphere as a whole are even smaller. To put them
 in context, note that despite the phase-in of regulations on
 CFC production during the rest of the decade, the chlorine
 content of the stratosphere will probably increase by a
 further ten to twenty percent. The ozone depletion predicted
 to follow from Ariane launches is likewise negligible compared
 with the impact of CFC emission. The ozone depletion averaged
 over the globe is significantly less than 0.1%.

 **> Conclusion

 The models used to predict the effect on ozone of various
 constituents depend on many assumptions that can only partly
 be verified, although much effort is spent over the world in
 substantiating them. The discrepancy factor between the
 prediction of models and actual measurements has decreased
 over the years and the trends are now well predicted. The
 qualitative confidence we can have in the calculations is
 therefore good.

 If we consider the other combustion products released by
 Ariane-5, we can discount the effects of the nitric oxides,
 which could appear only in very small quantities as a result
 of secondary reactions of the exhaust gases with ambient air.
 The main product released by the central core of Ariane-5 is
 hydrogen and water vapour. It can be expected that the
 question is reduced to that of the presence of water
 vapour. The effect of water generated by Ariane-5 should be
 less than that of chlorine and can probably be discounted. We
 shall however make certain of this by monitoring the results
 of the studies carried out by those organisations interested
 in hypersonic flight and those which deal with much larger
 amounts of water released in the atmosphere than Ariane-5
 could cause.

 Another unknown is the fact that the exhaust of our solid
 propellants contains particles of alumina of 80-300 micrometer
 diameter, as a combustion product. It is known that alumina is
 used in certain reactions as a catalyst and the question is
 open whether these particles can influence the ozone
 equilibrium. It is now known that surface catalysis plays an
 important role in producing the Antarctic ozone hole. The
 study of the impact of alumina is a possible subject for
 further study.

 Although chlorine emissions are unfavourable on a large scale,
 those due to launch activities are small even compared to
 natural sources. Curtailing space activities would bring
 advantages so small as to be non-measurable, but offset by
 major disadvantages and disruptions. The only recommendation
 that one can make is to remain open to new findings, new and
 updated results, so as to make sure that launch activities
 remain free from the charge of degrading the ozone layer.

 J.F. Lieberherr
 ESA


 ***> Post injection sequence of Ariane-4

 ===> It is well known that the spent Ariane first two stages
 fall back harmlessly into the atmosphere and into the sea.
 Since the H10 third stage goes into the final launcher orbit
 (geostationary transfer orbit: GEO, low Earth orbit: LEO; sun-
 synchronous orbit: SSO . . .) with the payload, very specific
 manoeuvres have to be performed in order to guarantee a
 harmless reentry and destruction of the stage, once its
 mission is completed.

 After H10 thrust termination, a complex sequence of
 orientation and spin manoeuvres is carried out to orient and
 spin the payloads according to mission requirements. These
 manoeuvres are performed by the H10 attitude and roll control
 system (SCAR: systeme de controle d'attitude et de roulis) fed
 by the pressurant gas (GH2) from the LH2 tank. After
 separation and in order to prevent any risk of H10/payload
 collision, the stage is moved away from the satellite
 (avoidance manoeuvre) by expelling pressurant gas (He/G02
 mixture) from the LOx tank through the GO2 venting nozzles.

 The sequence and the duration of these manoeuvres is
 programmed so that the difference in internal pressure between
 the LH2 and LOx tanks remains always positive; this is
 required to prevent damage to the common tank bulkhead.

 However, an explosion of a spent H10 stage has been observed
 by NORAD during routine space debris tracking, after flight
 V16 to sun-synchronous orbit. Debris in SSO are a concern
 because of the long time needed to allow the drag forces due
 to the traces of atmosphere, to slow the debris down and bring
 them to burn in the atmosphere. Possible debris in LEO are
 less of a concern because of their short lifetime.

 Analyses have allowed tracing the H10 explosion to the
 break-up of the LH2/LOx tank; this could result from the
 uncontrolled pressure build-up generated by the vaporisation
 of the residual propellants (typically 120 kg LOx and 150 kg
 LH2) and a simultaneous failure of the pressure relief valves
 to remain open.

 In order to prevent pollution of outer space and in particular
 of the LEO volume occupied by the present and future manned
 missions, it was decided to `passivate' the H10 3rd stage
 after completion of its mission; this consists in rendering
 the stage fully inert through an appropriate venting sequence
 during which the above-mentioned requirement on pressure
 levels (PH2 > P02) is respected. 

 A rather simple solution has been found. It consists in
 changing the operating mode of the S34/GO2 and S37/GH2
 pressure relief valves by means of added pyrotechnic devices: 
   - the S34 pyro rod is activated once the avoidance manoeuvre 
     is completed (end of launch mission) to definitively block 
     the pressure relief valve in open position; thus complete  
     emptying of the LOx tank will follow; 

   - the subsequent opening of the S37 pyro valves opens the    
     pilot compartment of the pressure relief valve to space    
     vacuum which has two consequences:

     *  first the S37 setting point (opening threshold) is      
        lowered from typically 3 bar to 0.8 bar;

     *  then it creates a by-pass to the pressure relief valve  
        when this one closes definitively (LH2 tank pressure    
        below 0.8 bar); thus a permanent venting allows         
        complete emptying of the LH2 tank.

 This passivation system has been successfully used for the
 first time on flight 35 (22 January 90); it was an Ariane 40
 mission in SSO with a single payload (Spot-2). Next
 application will take place with the same launcher version on
 a similar orbit (flight 44/ERS-1).

 The passivation scheme has been limited so far to this type of
 mission and launcher version for two main reasons:

   - because pollution of this space volume is specially        
     critical to manned flights;

   - because of the availability on AR40 of the four additional 
     pyrotechnic circuits (for the two added pyro systems)      
     which are occupied on the other Ariane versions (booster   
     pyrotechnic commands for example).

 Analyses have shown that the extension of the passivation
 measure to all Ariane versions and missions is technically
 feasible but with some unfavourable mass and cost impacts. The
 decision on whether to apply generally this technique will
 have to consider this commercial aspect.

 P. Luquet
 ESA