Manufacturing technology of the Romanian liquid rocket engine MRE-1B.

Rugescu, Radu Dan ; Aldea, Sorin ; Silivestru, Valentin 等

Abstract: The first Romanian (after Oberth) liquid propellant rocket engine MRE-1B was developed during the ADDA program at the University "Politehnica" of Bucharest. The 57th International Astronautical Congress in 2003 proved it the only East-European liquid propellant rocket system develop outside USSR after world war II. The specifically developed unit for static test firings on the test stand at normal atmospheric conditions includes original manufacturing solutions within a general design background similar to the Walter engines. The contribution to the technology stands in the dismantable design for all engine blocks, including the cooling jacket and the spray-injector head. The variable geometry version of the engine, currently into the patenting procedure, is announced.

Key words: Rocket engines, space propulsion, manufacturing.

1. MRE-1 DESIGN CONCEPT

At the level of the year 1960 little experimental knowledge on rocket propulsion existed in Romania, despite the large international literature and the existing soviet rocket armament on Romanian territory (Rugescu 2003). With a lot of challenge for the designers, natural precautions were engaged to minimize the risks throughout the research, justified in this risky field of technology. As a result, the following parameter levels and work conditions were chosen (Rugescu & Marculescu 1967):

1. Ground level steady thrust F=200N;

2. Chamber steady pressure level [p.sub.c]=7 bar;

3. Exit steady pressure level [p.sub.e]=1 bar;

4. Longest possible run time [t.sub.a]=1 min;

5. Non-cryogenic liquid propellant;

6. Low level propellant toxicity;

7. Non-self reacting (non-hypergolic) components;

8. Auxiliary fuel chemical ignition;

9. Remotely-controlled firing (70 m);

10. All-portable energy/water supplies.

These general conditions were found better satisfied by:

1. White-fuming nitric acid (96%) at [alpha]=0.62 excess ratio;

2. Jet-engine type T-1 kerosene as main fuel STAS 5639/57;

3. American "Aerobee" fuel (75/25 aniline and furfuril alcohol) as starting, hypergolic compound;

4. Volume-pump, controllable rate propellant supply system;

5. Low pressure, external water cooling of the entire motor;

6. Water purge of near-engine devices after each firing;

7. Lead, car-type, 24V d. c. electric power supply.

A low-pressure, volume-gear-pump main propellant feed system was embedded to accomplish the safety requirements. A minimal set of engine's block parameters was chosen for measurement, display and recording on the control panel:

1) Unsteady axial gas-dynamic thrust;

2) Pressure level in the thrust chamber.

The turning rate of pumps shaft for oxidizer and fuel, the main propellant delivery pressure at exit of pumps, the main propellant flow rate of components at engine entrance, the feed pressure for the auxiliary starting fuel, the ignition time delay and the cooling water temperature were all recorded. The capability of rocket engines manufacturing in the existing Romanian industrial environment was demonstrated.

2. TECHNOLOGY OF THE ENGINE BLOCK

The technological sketch for the engine block is shown in fig. 1. The cooling jacket around the most curved zone of the nozzle is solved into a single piece (item 12), due to the low expansion ratio. Its geometry ends in a very short conical nozzle indeed, with a theoretical ratio in exit-to-throat diameters E (Rugescu 2003) of E [equivalent to] [D.sub.e]/[D.sub.t] = 1.325 as seen in Fig. 1 and 3. With a general thickness of 2.5 mm of the walls a 3 mm cooling spill allows the nozzle (11) to glide freely into the single piece nozzle jacket (12), without any need for further sophistication. This highly simplifies the engine's block design.

An elastic element under each of the four nuts that seal the cooling channel is provided to balance the thermal dilatation.

[FIGURE 1 OMITTED]

[FIGURE 2 OMITTED]

3. MANUFACTURING TECHNOLOGY

Specific technologies are used for the manufacturing of the injector collector (5), injector head (2), thrust chamber (15) and nozzle (11, Fig. 1). All the details from the thrust chamber side of the collector are turned in the lathe up to finishing, within the quality of the sealing plane below the toroidal channels (Fig. 2).

The upper side is mostly a flat surface with four holes for the mounting screws and two threaded holes for the two inflow nipples. On the side area a flattening accommodates the threaded hole for the chemical ignition fuel. These are processed from a second machine gripping by means of two dedicated manufacturing devices.

The nozzle piece is cut by turning from a plane bar of INOXITERM-5 Ni-Cr refractory alloy (Rugescu 2003). First the entire inner surface is threaded with two fastenings upon the external cylindrical rough surface, when half of the external surface is also processed. The one-shot processing of the interior allows secured coaxiality and diameter tolerances in the supersonic part of the effuser. The remainder of the outside surface, with no major size or geometry restrictions, is further cut while a device, similar to the thrust chamber and screwed up in the M56x1 thread, fastens the piece in the lathe.

The injector head (Fig. 4) is turned on the entire surface from a 1H18N9T, 18/8 stainless steel bar into a single machine fastening, from the upper side. The central hole for pressure pick-up is also practiced in this stage on the thread. The upper 4 mm prominence of 5.5 mm diameter is cut the last. The cylindrical hollow of 4 mm in diameter, used to fasten the head into the correct position through a mounting pin during assembly, is lastly drilled by milling. On a coordinates-drilling machine the fine, very low diameter injector holes are finally processed with a regular fastening device.

[FIGURE 3 OMITTED]

[FIGURE 4 OMITTED]

[FIGURE 5 OMITTED]

The thrust chamber (15) is considered in three versions with different length to allow searching the effect of the chamber volume upon the thrust efficiency Is. At a diameter of 52 mm with its 2.5 mm wall thickness the short (150 mm) and medium (360 mm) chambers put no problem and a single fastening into the thread allows the whole inner/outer cutting and finishing.

For the long version of the chamber (510 mm) its nozzle end is supported into a usual rotating cone on the lathe to accomplish the rigidity requirements. This manner any observable vibration was avoided during the finishing phase.

4. CONCLUSIONS

A simple and safe technology had well qualified to manufacture the Romanian liquid rocket engine MRE-1B, as given above. It contributed to the success of the first static firing of the engine.

5. REFERENCES

Rugescu, R. D. (1999), ADDA rocket engines research program: principles and thermo-chemical data, Sci. Bull. Univ. "Politehnica" Bucharest, series D Mech. Eng. 61, 3-4(1999), pp. 405-414

Rugescu, R. D. & Marculescu, R. (1967), Design and Manufacturing of the MRE-1B Rocket Engine and of its Test Stand, National Student Conference, Timisoara Romania, November 1967

Feodosiev, V. I. (1963), Calculus of strength for thermally loaded parts of liquid propellant rocket engines, Mashinostrojenie, Moscow

Messerschmid, E. & Fasoulas, S. (2000), Raumfahrt-systeme, Springer Verlag, Berlin N.York

Rugescu, R. D. (2003), Beginnings of Space Propulsion Research in Romania, Paper IAC-03-IAA.2.3.04, Proceedings of 54th International Astronautical Congress, 29 Sep-03 Oct 2003/Bremen, Germany.