TOWARDS A HEAVY LAUNCHER - PROPULSION SOLUTIONS - A. Souchier - C. Rothmund Snecma Moteurs, Direction Grosse Propulsion à Liquides

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1 Souchier_2002 TOWARDS A HEAVY LAUNCHER - PROPULSION SOLUTIONS - A. Souchier - C. Rothmund Snecma Moteurs, Direction Grosse Propulsion à Liquides ABSTRACT The Martian human missions will need heavy launchers with payloads in earth orbit in the range of 80 to 150 tons. Previous launchers of this size were the American Saturn 5 and the Russian Energya. The industrial tool to produce these launchers is no more available. Some projects more or less based upon the shuttle hardware have been studied in the past years in the USA (Shuttle C, Magnum, Star Eagle, ). In the beginning of the 90s, in Europe, CNES studied a heavy version of ARIANE 5. The development of a heavy launcher will probably be, as the whole Martian program, an international endeavour building on the engines and stages operationally used for commercial launchers in the next twenty years to reduce the costs. Thus we may already have an idea of the configuration of this heavy international launcher. The need for such a launcher will probably remain unchanged for decades. For Martian missions the progresses in space propulsion will reduce the propellant mass and/or reduce the travel time but the low earth orbit mass will remain the same. Also the 80 or 150 tons low earth orbit payload will fulfil the needs for lunar and other planets than Mars exploration as well as eventual heavy commercial platform in geo-stationary orbit. INTRODUCTION The existing large size launchers are more or less able to send 25 tons in low earth orbit (LEO). ARIANE 5 and the US Space Shuttle have this capacity. Proton M and Titan 4B have, in those low earth orbits, a payload slightly above 20 tons. But for ambitious missions to Mars or to the moon, particularly manned missions, a payload between 4 and 8 times those values is necessary. THE PAYLOAD NEEDS The calculated low earth orbit payload needed for a manned Mars mission has evolved since 50 years from huge extraordinary values to more reachable numbers. In 1949 Wernher Von Braun designed a manned Mars mission starting from tons in earth orbit! The mission objectives were ambitious : 70 astronauts were to be sent to Mars. With a more classical crew size around 7 people the payload could have been reduced by an order of magnitude. At the end of the sixties and beginning of the seventies new projects are studied in the US. The low earth orbit payloads are ranging from 1500 to 2800 tons. Launchers with cryogenic hydrogen and oxygen upper stages and also escape stage using nuclear thermal propulsion, have considerably reduced the LEO payload needs. In the eighties a new idea arises : the use of the Martian atmosphere to brake the incoming vehicle and insert it without propellant consumption in a low Martian orbit. The process is not completely free because the vehicle has to be protected by a thermal heat shield and also a small amount of propellant has to be used for orbit adjustment. The heat shield mass may be lower than 15 % of the Page 1

2 vehicle mass. This value has to be compared to 30 to 60 % of the vehicle mass devoted to in orbit insertion propellants in case of a pure propulsive braking. Roughly the aero-braking in orbit insertion divides by 1.5 the initial LEO payload needs. Later on, at the beginning of the nineties, Robert Zubrin and David Baker then working for Martin Marietta, proposed to manufacture propellants from in situ resources on Mars to fill the earth return vehicle tanks. In the new scenario less than 6 % of the return propellant is brought from earth. This mass is liquid hydrogen which is converted on Mars by reaction with the atmospheric carbon dioxide (CO2) in methane (CH4) and oxygen. Other means of producing propellant on Mars will also be available : water electrolysis for example. Thus a new reduction step is reached : the initial mass in low earth orbit may be again divided by years of Martian projects A sharp decrease in LEO mass : -Higher propellant efficiency (hydrogen, oxygen) -Mars atmosphere for aerobraking -In situ propellant production NASA For a human mission the necessary LEO mass seems now to be in the order of 300 tons in general splitted in two vehicles, one automated and one manned. If a redundancy is wished (which seems necessary) then three 150 tons vehicles have to be sent in earth orbit. Page 2

3 It is interesting to note that for manned lunar missions the 150 tons value is also the right need. It was roughly the Apollo Saturn parking orbit mass (~ 140 tons). In the latest NASA Design Reference Mission (DRM4) the escape stage uses nuclear thermal propulsion. The proposed LEO masses for the vehicles are 135 tons for the automated and 168 tons for the manned. The manned vehicle does not aero-brake upon arrival at Mars but is inserted in Martian orbit by a retro-propulsive burn from the nuclear thermal propulsion engine. This solution has been chosen for safety considerations mainly because an aborted mission (return to earth) without landing on Mars remains possible from Mars orbit. In this Design Reference Mission the automated vehicles are launched from earth in two parts. The final vehicle is assembled by in orbit rendezvous. For the manned vehicle there is also two heavy launchings plus one shuttle flight for the habitat and crew transfer. Thus 150 tons remain the mass to be launched for the basic spaceship. In the DRM 4 mission the authors have chosen two launchings to build the basic spaceship from two 70 tons payloads. For the whole mission 6 heavy launchers and one shuttle flight is needed. 4 rendezvous are conducted. With a 150 tons heavy launcher capacity, those numbers would be reduced to 3 heavy launchers and one shuttle flight and only 1 rendezvous. Such a mission build up would be far more reliable. FUTURE EVOLUTION OF THE PAYLOAD NEEDS Compared to hydrogen/oxygen liquid propulsion, the use of nuclear thermal propulsion with a specific impulse nearly two times higher should bring a reduction in the LEO mass to be orbited for a given mass sent to Mars. Unfortunately nuclear thermal engines which include a nuclear reactor and radiation shields are heavier than hydrogen/oxygen engines. They also use hydrogen as ejected fluid (which is better for specific impulse owing to its low molecular mass). Hydrogen has a very low density of 70 kg/m 3 and the tank mass is high for a low hydrogen mass content. Thus the nuclear thermal stage dry mass is high. For a minimum energy travel to Mars, a 50 tons payload will require 66 tons of hydrogen and oxygen in a 6 tons dry mass stage with classical rocket propulsion. With nuclear thermal propulsion the dry mass is about 25 tons for 41 tons of hydrogen. Thus the benefit of nuclear thermal propulsion is only 10 % in LEO mass reduction. If we shift to nuclear electric propulsion, there is still a mass penalty which gives stage dry masses in the vicinity of 25 tons. But the electric propulsion higher specific impulse (1 500 s for plasmic propulsion) reduces the propellant mass to 24 tons for the same 50 tons sent to Mars. The total LEO mass is reduced by 20 % compared to liquid propulsion. This result is obtained after taking into account that, with the low thrust linked to electric propulsion, a 20 % penalty has to be paid in DV for gravity losses on the trajectory. If the specific impulse is increased to s, as possible for various type of electric propulsion, there may be again a 20 % reduction in LEO mass. Compared to chemical propulsion the LEO mass is reduced from 125 tons to 78 tons for low energy trajectories to Mars Page 3

4 . The differences between all these propulsion modes are increased when the DV needs are higher if short trip, high energy missions are targeted. For 5 km/s DV, the LEO mass for 50 tons to Mars is 180 tons in chemical H2/O2 propulsion, 110 tons in plasmic propulsion (ISP at s) and 80 tons in electric high ISP ( s) propulsion. The LEO masses are given in the chart below depending on the propulsion mode and the speed to be given from low earth orbit. With this speed 2 typical Earth-Mars transition travel duration are given : one when Mars is at the smallest distance from the sun and the other one when this distance reaches its maximum. Thus the LEO payload requirement for a heavy launcher may be reduced to 80 tons for future needs. This is also compatible in a first period with 160 tons vehicles built after a rendez-vous. But a 150 tons payload class vehicle will authorise classical chemical propulsion missions as well as future high speed mission with plasmic propulsion in one single launch. In summary, a good requirement for a heavy launcher seems to be the range 80 to 150 tons. HEAVY LAUNCHERS STORY One of the first serious heavy launcher project was Von Braun's "Ferry Rocket" presented in the 1952 "Mars Project". Using hydrazine and nitric acid as propellants the rocket has a T mass at Page 4

5 lift off for a moderate 39 tons payload in low earth orbit. The 3 stages are recovered and reused. The first stage is powered by 51 engines developing a total thrust of T. Fifteen years later the Apollo program Saturn 5 launcher reaches 3.6 times the "Ferry Rocket" payload for only 0,46 times its lift off mass! Better propellants (liquid oxygen and kerosene on the 1st stage but mainly liquid hydrogen and oxygen in the upper stages), better dry masses for the stages (but the stages are not recoverable) have brought an impressive improvement on the preceeding design. NASA With 140 tons in LEO orbit, a performance reached nearly 40 years ago, the Saturn 5 demonstrates the feasibility of the heavy launcher needed for the 21st century ambitious missions! In the same time frame between 1969 and 1972 the Russians flew 4 times a heavy launcher, the N1. This launcher was also designed for a manned moon trip. It did not use cryogenic propellants but partially compensated by using liquid oxygen / kerosene high specific impulse engines. The absolute and relative performance was slightly lower than the Saturn 5 with 95 tons in LEO for a tons mass at lift off. None of the 4 flights succeeded. Page 5

6 Afterwards the Russians developed a new heavy launcher Energya which was intended for a dual use : heavy payloads or a space shuttle. The first flight occurred in May 1987 and the second with the Buran space shuttle in November Then with economical difficulties rising in the soviet union the program was cancelled. Energya was propelled at lift off by four kerosene/oxygen boosters and a central core using hydrogen and oxygen. The boosters equipped with a RD tons thrust engine are still in use as the Zenith launcher first stage. The central core was powered by 4 RD O120 engines 200 tons thrust each. The payload reached 90 to 100 tons for tons at lift off which gives a ratio slightly lower than the Saturn 5 (4,0 % compared to 4,75 %). Heavy version of Energya with 6 and even 8 boosters were foreseen to launch 150 and 200 tons in low earth orbit. In April 1981 the American space shuttle flew for the first time. Propulsion options chosen were and still are 2 solid boosters to get a high thrust at lift off and a cryogenic hydrogen and oxygen propulsion in the central core to get a high specific impulse afterwards. The solids develop tons thrust each and the 3 cryogenic SSME engines 215 tons in vacuum. The lift off mass is around tons for a payload in the shuttle cargo bay of 25 tons. But the whole shuttle outside the propulsion bay is the payload. Page 6

7 Seen as a heavy launcher the space shuttle is a 80 tons class vehicle. NASA DESIGN RULES FOR A FUTURE HEAVY LAUNCHER The needs for heavy Martian or lunar missions seems to arise somewhere after The heavy launcher development should start around It is supposed that new specific developments will be reduced to what is absolutely needed for such a launcher and that, as much as possible, available systems, components, equipments will be taped off from existing launchers or existing developments for medium size commercial launchers. To reduce operational costs it is also interesting to use the systems, components, equipments which will be in production for a long time (> 2030) for those operational commercial launchers. Thus the industrial production basis and organisations, the teams, will be used with the best efficiency reducing the costs. It is also known that large production rates increase the hardware reliability because the teams have a large experience and a lot of knowledge is brought by the hardware flight operations. Also the future heavy payload missions will be international missions. A future heavy launcher development will surely be international and the launcher main systems will come from the different countries or group of countries involved in these heavy missions. Thus possible configurations of those future heavy launchers may be defined. SHUTTLE DERIVATIVES Using shuttle derivatives as heavy launchers is a logical way to go in USA. Launchers like Shuttle C, Shuttle Z, Ares have a LEO capacity ranging from 80 to 120 tons. On the Shuttle C (for cargo) only the propulsion bay is kept from the orbiter. Above the propulsion bay, the cargo bay is replaced by a classical fairing. Page 7

8 NASA The Shuttle Z may be seen itself as a Shuttle C derivative. The propulsion bay is fitted with 4 SSME engines. The fairing above is far larger and includes a cryogenic hydrogen / oxygen upper stage with 127 tons of propellant. This stage is powered by a simple SSME or a cluster of smaller engines. The Ares is also a shuttle derivative in which the payload fairing is located above the external tank to reduce drag. A four SSME engines propulsion bay is located at the lower part of the stack on the side, as in the orbiter propulsion bay on the baseline shuttle. Ares has a capacity of 105 to 115 tons in LEO or is able to send 40 tons in a earth - Mars trajectory. In all cases, the final boost is given by an upper stage loaded with 160 tons of hydrogen and oxygen and carried under the fairing. Owing to the use of shuttle components those launchers are not opened for a possible international co-operation excepted for the upper stage. The Magnum heavy launcher makes use of a mix of components coming from the Space Shuttle and expendable launchers being developed. The NASA current martians Design Reference Missions are built on launches with a Magnum vehicle. The core stage is derived from the shuttle external tank equipped with a 2 RS 68 engines propulsion bay. The RS 68 is a powerful 290 tons thrust engine developed for the Delta IV ELV launcher. The boosters may be either shuttle SRM boosters or future liquid fly back boosters yet to be developed for the current shuttle. An upper circularization stage is derived from the Delta III or IV upper stage. The payload would reach till 90 tons in LEO. Page 8

9 The liquid fly back boosters may be open to international co-operation. They use as a baseline the RD 180 liquid oxygen/kerosene mission engine (8 engines rated at 380 tons thrust each). ARIANE 5 DERIVATIVES Starting from 25 tons in LEO it may seem difficult to extrapolate a 100 tons class vehicle from the present ARIANE 5. However in 1991 CNES, the French national space agency, conducted some studies to show that, using ARIANE 5 components and systems, a heavy launcher could be developed. The proposed heavy ARIANE 5 would use 4 solid boosters compared to 2. The central core stage would be new but would be fitted with 5 VULCAIN 2 engines compared to one on a regular ARIANE 5. The upper stage would be derived from the present ARIANE 5 EPC central core (same diameter and propulsion system but shorter tanks). ARIANE 5 HEAVY and ARIANE 5 The new central core would carry 620 tons of hydrogen and oxygen. The upper shortened EPC would be de-rated at 70 tons compared to the present 170 tons. Total mass at lift off is 1874 tons. The LEO capacity reaches 80 tons without using the upper stage. A 35 tons payload may be sent to the moon and a 26 tons payload to Mars. This launcher would be able of Apollo class missions. The payload to the moon is smaller than a Saturn V but with use of hydrogen and oxygen to land on the moon (compared to storable for Apollo) the useful payload on the moon would be the same. To illustrate with a smile this Apollo capability some of the heavy ARIANE 5 drawings are shown with an escape rocket and Apollo type capsule in the upper part. Page 9

10 As presented in this paper first part, 80 tons is the lower limit for a heavy launcher and some studies were conducted to upgrade this payload. Replacing two ARIANE 5 EAP solid boosters by two shuttle SRB the payload is boosted to 100 tons. A heavier stack may be built with 4 SRB s and 2 EAP s. The 4 SRB s are ignited with the 5 VULCAIN central core at lift off. After two minutes the SRB s are jettisoned and the EAP s ignited. The payload is then boosted at around 110 tons. This family with the same central core loaded at 620 tons of hydrogen and oxygen offers a range between 80 and 110 tons depending on the boosters used. The payload may still be increased by stretching the central core. At a 745 tons propellant load, the payload reaches 130 tons. At 920 tons the payload is then at 140 tons. With its new central core stage at 8,2 m in diameter which is the shuttle external tank diameter, and the use of SRB s this launcher development could be a European - US endeavour. Naturally Russian boosters could also be considered. Fundamentally this booster is not intended to launch astronauts but the heavy payloads needed for human lunar or Martian missions. The astronauts are supposed to be launched by a human rated vehicle like the shuttle and rendezvous in LEO with the spaceship orbited by the heavy launcher. The heavy launcher could be used also for future heavy scientific missions and heavy geo-stationary platforms. In all those missions reliability is the prime objective. In all the Martian trip scenarios studied to day, the failure for one of the heavy payloads to reach LEO would mean the mission cancellation. With the use of solid boosters the proposed design offers a high reliability. The reliability (ability to fulfill the nominal mission) is recognised one order of magnitude higher for solids than for liquids. Safety for eventual astronauts is lower than for liquid because when a solid booster failure occurs it is far less manageable than for a liquid booster failure when in general the engines may be stopped before catastrophic consequences. For the central core with cryogenic propulsion which is needed for the mission performance, the 5 VULCAIN configuration offers an engine out capability such that the mission may be fulfilled. A configuration with only two engines like the RS 68 would not bring the same capability. All the payloads calculated for this heavy launcher family are given with ARIANE 5 Evolution type hardware (VULCAIN, boosters) to demonstrate that the capacity already exists. With foreseen hardware for ARIANE 2010 new generation of ARIANE, very large improvements will be obtained. Also now the upper escape stage may use the VINCI engine which did not exist at the beginning of those heavy launchers studies in Page 10

11 CONCLUSION In this decade all the main building blocks to develop a heavy launcher are available. The situation is far better than when the USA developed the Saturn 5 or the USSR the N1. With the lessons learned from international co-operation (space station and others) a co-operative development will also be more easily manageable. Page 11