ROCKET ENGINE, ROCKET AND START METHOD OF ROCKET ENGINE Technical Field The present invention relates to a rocket engine, especially, to a liquid rocket engine using hydrocarbon as fuel. Background Art [000/] A Liquid rocket engine using hydrocarbon as propellant is known. The liquid rocket engine has a turbo pump for supplying the propellant (oxidizer and hydrocarbon). The turbo pump is driven with the rotation of a turbine. The turbine is rotated with a combustion gas which is generated by an auxiliary combustor called a gas generator (sub combustor) which uses hydrocarbon as fuel, and LOX (Liquid Oxyqen) as ox:idLo:er. For example, Non-Patent Literature 1. discloses the liquid rocket engine to which such a gas qenerator cycle is applied. [0003] FIG. 1 is a diagram schematically showinq Lhe configurations of the liquid rocket engines described in Non-Patent Literature 1. In FIG. 1, a diagram in the upper Jeft portion shows the liquid rocket enqinc to which a gas generator cycle is applied. In th~s rocket engine, a fuel pump and an oxidizer pump arc used as turbo pumps. The fuel pump is driven by a f u e J L u r b i n e , and the ox i d i z e r p urn p i s d r i v en by an oxidizer the turbine. Thus, the fuel turbine and the oxidizer turbine are driven by the combustion gas generated in the gas generator. [00011 When such a gas generator cycle is applied to -~x - the liquid cocket engine, there are the following problems. The first problem is in that an additional combustor becomes necessary. Because the additional combustor becomes necessary, there is a fear that the reliability decreases and a cost becomes high. The second problem is in that the combustion gas used to drive the turbines is exhausted just as it is. Because the energy of the combustion gas used to drive the turbines is not effectively utilized, an energy loss is big. The third problem is in that a very high reliability is required for the gas generator. Though the operation environment is severe because the gas generator is used for the combustion, a trouble of the gas generator is fatal to the rocket engine. When the gas generator breaks down, the liquid rocket eng~ne loses a turbine driving force immediately so that it becomes difficult to supply the fuel and the oxidizer. In this way, the system of the rocket engine is not robust. An upper right portion of FIG. 1 shows a rocket engine using a staged combustion cycle. In this rocket engine, a precombustor is provided in front of the combustor, and a precombustion gas of the precombustor drives the fuel turbine and the oxidizer turbine and is supplied to the combustor. In this case, because the precombustion gas after driving the turbines ls further used for the combustion, the above second problem does not occur. However, because the precombustor as an auxiliary combustor is used, tho first and third problems remain. Also, a lower portion of l·'IG. 1 shows the rocket engine using an expander cycle which uses vaporization and expansion of the liquid-hydrogen. It is difficult to apply this rocket engine to the rocket engine using hydrocarbon as the-; fuel. Citation List [Non-PatcnL Ijitcrature] [0006] [Non-Patent Literature 1] George P. Sutton, Oscar Biblarz, "Pocket Propulsion Elements", Seventh Edition, John Wiley & Sons, Inc., p. 223 (Figures 6 - 9), (;700}). Summary of the Invention [ 0 0 0 7 l 'L'hcrE;fore, an object of the present invention is to provide a rocket engine which has a high reliability and few energy losses and which uses a hydrocarbon fuel. Also, another object of the present invention is to provide a robust rocket engine using a hydrocarbon fuel. Also, still another object of the present invention is to provide a rocket engine in which the mixing of fuel and oxidizer is promoted, in which the combustion efficiency is improved and the exhaust of gas is made unnecessary, and which uses the hydrocarbon fuel. I o o o tl I Thes<::"~ objects of the present invention, and objects and advantages except for them can be easi1y confirmed by the following description and the attached drawings. [00091 A rocket engine of the present invention is provided with a fuel passage of pipes, a catalyst section, a turbine, a first pump, a combustion chamber and a nozzle. A hydrocarbon fuel flows through the fuel passage of pipes. The catalyst section is provided on the way of the fuel passage of pipes to gasify the fuel. The turbine is provided on the way of the fue1 passage of pipes and is driven with the gasitLed fuel. The first pump supplies the fuel to the fuel passage through the drive by the turbine. The combustion chamber combusts the gasified fuel supplied from the fuel passage of pipes and the oxidizer. The nozzle sends out the combustion gas from the combustion chamber, and carries out heat exchange with a part of the fuel passage of pipes for the nozzle to be cooled. [ 0 0 1 0 l In the above-mentioned rocket engine, the fuel passage of pipes may have a first path and a second path. The first path leads the fuel to the combustion chamber after the heat exchange with the nozzle and gasification of the fuel in the catalyst section. The second path leads the fuel to the combustion chamber just as it is. The combustion chambec may combust the mixed fuel of the gasified fuel supplied through the first path and the fuel supplied through the second path. [00111 In the above-mentioned rocket engine, the catalyst section may be provided inside the fuel passaqc of pipes at a location where the heat exchange with the nozzle is carried out on the way of the fuel passage of pipes. [001~] rn the above-mentioned rocket engine, the catalyst section carries out thermochemical decomposition of the fuel to gasify the fuel. The endothermic reaction of the thermochemical decomposition may cool the nozzle. [ 0 0 l 3 1 ln the above-mentioned rocket engine, the catalyst section may be formed in a layer to cover the inner wal 1 of the fuel passage of pipes. [0014] rn the above-mentioned rocket engine, the catalyst section may be provided on a rear side of a location which is on the way of the fuel passage of pipes and in which the heat exchange with the nozzle is carried out. In the above-mentioned rocket engine, the catalyst section may be formed from a plurality of catalyst grains in a container. [0016] The above-mentioned rocket engine may furLher include an oxidizer passage of pipes and a second pump. The ox_idizer flows to the combustion chamber through Lhe oxidizer passage of pipes. The second pump supplies the oxidizer to the oxidizer passage of pipes through the drive by the turbine. The turbine, the first pump and the second pump are connected with the same rotation axis. [00171 A rocket of the present invention is provided with the rocket engine, a fuel tank and an oxidizer Lank. The rocket engine is described in the above. The fuel tank is connected with the fuel passage of the rocket engine. The oxidizer tank is connected with the oxidizer passage of the rocket engine. [OOUll A sLart method of a rocket engine according lo the present invention includes supplying a hydrocarbon fuel and an oxidizer to a combustion chamber by using a pressure to combust there. Moreover, when the combustion chamber is heated through the combustion, the method includes supplying the fueJ to Lhc catalyst section by using the pressure while cooling a nozzle by heat exchange, to gasify the fuc L. [0019] Moreover, the method includes driving Lhc I -ftturbine: wi Lh the qas_ified fuel. Moreover, the method includes driving the first pump by the turbine to gasify a part: of the fuel in a catalyst section while cooling the nozzle through the heat exchange. Moreover, the method includes driving the first pump by the turbine to supply a remaining part of the fuel to the combustion chamber. Moreover, the method includes driving a second pump by the turbine to supply the oxidizer to the combustion chamber. Moreover, the method includes combusting the gasified fuc l, Lhe supplied fuel and the supplied oxidizer in the combustion chamber. [00/.0l According to the present invention, in the rocket: engine using the hydrocarbon fuel, the rc;l iabi l i t:y can be made higher, and the energy loss can be made few. Also, in the rocket engine using the hydrocarbon fuel, the robust property can be improved. Also, in the rocket engine using the hydrocarbon fuel, the mixing of the fuel and the oxidizer is promoted, so that the exhaust of gas is eliminated and the combustion cffLciency can be improved. Brief Description of the Drawings [00/.1] FTC. 1 is a diagram schematically showing Lhc configuration of a liquid rocket engine disclosed in Non-Patent 1, L Lerature 1. I<'TC. / is a diagram schematically showing Lhc configuration of a rocket engine according Lo a f rsL embodiment and a rocket to which the same is appJ i cd. FlG. 3 is a diagram schematically showing an example of a catalyst section of the rocket engine according to Lhe first embodiment. FiG. 1 is a flow chart showing a start operation of the rocket engine according to the r rsL embodLmcnt. s- y- FIG. 5 is a diagram schematically showing lhe configuration of the rocket engine according to a second embodiment and the rocket to which the same is appLied. Description of Embodiments [0022] l!c~reinafter, a rocket engine according to embodiments of the present invention and a rocket to which the rocket engine is applied will be described with reference to the attached drawings. [0023] [First Embodiment] The configurations of the rocket engine according to a first embodiment and the rocket applied with the same will be described. FIG. 2 is a diagram schematically showing the configurations of the rocket engine according to the present embodiment and the rockc~t appJ Led with the same. The rocket 1 flies by combusting the fuel of hydrocarbon by using the ox i d i zc r and by spouting out the combustion gas. The rocket 1 is provided with a rocket main unit 2 and a rocket cnqinc 3. [00/1\] The rocket main unit 2 has a fuel tank 1\, an oxidizer tank 5 and a control unit 6. The fuel tank fl, stores a Jiquid hydrocarbon fuel. The fuel tank 1sends out the fuel to the rocket engine 3 by means of qas pressurization and so on, based on the control of the conLro] unit 6. The oxidizer tank 5 stores a liquid oxidizer (e.g. LOX) The oxidizer tank 5 sends out an oxidizer to the rocket engine 3 by means of the gas prc;ssurJzation and so on, based on the control 0 f the control unit 6. The control unit 6 is an information processing apparatus such as a computer which has a pr-ocessor (e.g. CPU), a storage unit (e.g. RAM and ROM), and an interface, which are not illustrated. The control unit 6 may further have an input unit (e.g. a keyboard) and an output unit (e.g. a display). The control unit 6 controls the fuel tank 4, the oxidizer tank 5 and the rocket engine 3 at least through the interface. [0025] The rocket engine 3 is supplied with the liquid hydrocarbon fuel from the fuel tank 4 and with the oxidizer from the oxidizer tank 5, and combusts the fuel with the oxidizer to generate and spout out a combustjon gas. The rocket engine 3 includes a fuel passage of pipes 31 to 39, a catalyst section 51, a turbine 12, a first pump 11, a second pump 13, a combustion chamber 21, a nozzle 22, an oxidiz<:;r passage of pipes 41 to 43, a valve V1, a valve V2 and a r-otation axis 14. [ 0 0 2 6] The fuel passage of pipes 31 to 39 is a passage through which the hydrocarbon fuel flows. 'l'hc catalyst sectLon 51 is provided on the way of the fuel passage or pipes 31 to 39 to gasify the fuel. 'rhe turbine 12 is provided on the way of the fuel passage of pipes 31 to 39 and is driven with the gasified fuel. The first pump 11 is driven by the turbine 12 to supply the fuel to the fuel passage of pipes 31 to 39. 'rhe second pump 13 is driven by the turbine 12 to supply the oxidizer to the oxidizer passage of pipes 41 to 43. The first pump 11 and the second pump 13 ar -~- manufacture is provided as the catalyst section, Lhe increase of pressure loss can be suppressed. Moreover, Lhe operation temperature becomes low in Lhe catalyst section ~2 compared with the combustion c h a mb e r i n w h i c h a p r o p e ll an t i_ s c o mb u s t e d , s o t h a t Lhe operation environment is eased. Thus, the phenomenon such as a caulking can be further prc;venLed. is noL rL could be seen that the present invention tmiLed La each of the above embodiments and LhaL each embodiment can be appropriately modified or chanqed in a ranqe of the gist of the present invention. [OOS!\] The present invention is based on Japan Patent J\ppl ication No. JP 2013-030410 and clams a priority of it. The disclosure thereof is incorporated herein by reference. CLAIMS 1. A rocket engine comprising: a fuel passage through which a hydrocarbon fuel fLows; a catalyst section provided on a way of the fuel passage to gasify the fuel; a turbine provided on the way of the fuel passage and driven with the gasified fuel; a first pump configured to supply the fuel to the fuel passage through a drive by the turbine; a combustion chamber configured to combust the gasified fuel supplied through the fuel passage by using an oxLdizer; and a nozzle configured to send out the combustion gas and carry out heat exchange with a part of the fuel passage to be cooled. /. The~ rocket engine according to claim 1, wherein the fuel passage comprises: a first path configured to lead the fuel to the combustion chamber after the heat exchange and the qasification of the fuel in the catalyst section; and a second path configured to lead the fuel to the combustion chamber just as it is, wherein the combustion chamber combusts a mixture of the gasified fuel supplied through the first path and the fuel supplied through the second path. 3. The rocket engine according to claim 1 or /, wherein the catalyst section is provided inside the fuel passaqe in a location where the heat exchange ts carried out. ll . ·rr1c rocket enq· i. ne a ceo rd.ing to c l c1 _t m 3, wherein Lhe catalyst section gasifies the fuel through a thcrmochcmLcal decomposition with catalyst, and cools the nozzle with endothermic reaction of the thermochemical decomposition. 5. The rocket engine according to claim 3 or t1,, wherein the catalyst section is formed in a layer to cover an inner wall of the fuel passage. 6. The rocket engine according to claim 1 or 2, wherein the catalyst section is provided in a location where is on a way of the fuel passage and where is rear a heat exchange position with the nozzle. "/. The rocket engine according to claim 6, wherein the catalyst section comprises a plurality of catalyst grains in a container. 8 . The rocket engine according to any one of cLaims 1 to '/, further comprises: an oxidizer passage configured to supply an oxidLzer to the combustion chamber; and a second pump configured to supply the oxidizer to the oxidizer passage through a drive by thE; Lut"bine, whet"ein the turbine, the first pump and the second pump are connected with the same rotation axis. 9. A rocket comprising: the rocket engine according to any one of claims 1 to 8; a fuel tank connected with the fuel passage of the rocket engine; and an oxidizer tank connected with the oxidizer passage of the rocket engine. 10. A start method of a rocket engine, comprisinq: supplying a hydrocarbon fuel and an oxidizer to a combustion chamber using a pressure to combust therein; supplying, when the combustion chamber is heated through the combustion, the fuel to a catalyst section by uslng the pressure to gasify the fuel, while cooling the nozzle through heat exchange; driving a turbine with the gasified fuel; driving a first pump by the turbine to gasify a part of the fuel in the catalyst section while cool i nq the nozzle by the heat exchange; drivinq the first pump by the turbine to supply a remaining portion of the fuel to the combustion chamber; driving a second pump by the turbine to supply the oxLdizer to the combustion chamber; and combusting the gasified fuel, the supplied fuel and the supplied oxidizer in the combustion chambc r.