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The alternate solution offered by 10 Department of TsAGI, was reduced to the project of so-called ‘Siamese-Twins Tank’, which had three parallel cylindrical sections. Practically flat bottom surface of such external tank created a larger lift as a lifting body vehicle. It was offered to place the orbiter’s nose into a niche in the rear part of the tank. In this case the OPET had a very good aerodynamic characteristics and flight performance. The comparative analysis of launch trajectories of different OPET versions has shown that the version with ‘Siamese-Twins Tank’ had minimum total losses of characteristic velocity, and accordingly the maximum injected mass. Nevertheless, the proposal of TsAGI was not accepted, since the structural mass of such tank appeared much above any alternative ET versions.

Launch characteristics comparison for the RSTS of different types Table 4

It was necessary to find the version with the minimal mass of external tank, being the sole expendable element of the system. The solution offered by the Deputy General Designer Ya.I. Seletaky was very simple: to place oxygen in the rear part of ET. In this case the center-of-gravity position displaced back and the orbiter’s outer wing consoles of OP could be declined practically to a full angle without introduction of destabilizing wing at the ET nose. The mass of external tank was reduced at such layout because of decrease in the bending moment, acting on the tank’s construction. For the increase of the tank ability to withstand against external aerodynamic loads, the increased pressure inside ET was stipulated. These solutions, alongside with the application of aluminum-lithium alloys, give us opportunity to achieve low values of the relation of tank dry mass to the mass of fuel.

Many efforts were spent to find an optimal relative positional of orbiter and external tank within integrated structure of the MAKS 2-nd stage. There is an essential disadvantage of the basic concept with orbiter’s arrangement above the conical rear part of the external tank. In this case the orbiter’s bottom surface and wing are arranged under a large angle to the ET centerline, which decreases the lift-to-drag ratio of the MAKS 2-nd stage (OPET). More over, in this case a large angle of LRE thrust vector inclination downwards relative to the ET axis unfavorably influences the energy expenditure and loads exercised on OPET at the launch.

With the purpose to find the way to eliminate the indicated disadvantages, the parametric calculations with variation of an attachment point of OP nose on ET surface were conducted. For acceleration of work, the program module of OPET aerodynamic characteristics’ calculation was connected to STARTV program of trajectory calculation.

The results of researches have shown that moving OP along the upper surface of ET does not give any noticeable improvement of characteristics. The optimum by aerodynamics OPET layout was obtained at arrangement of the OP nose below under the rear part of ET. The second stage in this case reminds a barge with the towing vehicle-thruster (among aerodynamics and ballistics specialists the scheme has received a name ‘Mole’). The gain in main flight performance and aerodynamic loads of such scheme was amazing. Due to the increase of lift-to-drag ratio and optimal direction of LRE thrust along the ET axis the payload mass increase of approximately by 1.5 tons was achieved. In such scheme, the balancing characteristics and characteristics of stability and controllability were improved.

Despite of these impressive advantages, the OPET of ‘Mole’ scheme was rejected, though it was registered as invention.

The reason was that while considerably improving the characteristics of nominal flight, the ‘Mole’ scheme has resulted in complicating many processes in abnormal situations: more complicated has appeared the emergency separation of OP and ET, serious difficulties arose in designing the system of ejected seats for the crew. There were other disadvantages, in particular, approximation of LRE jet to the surface of the carrier-plane. Because of these problems it was necessary to refuse from the aerodynamically optimal layout of the ‘Mole’ OPET. As a result of overall analysis of different factors, the former scheme of MAKS second stage with the upper arrangement of the orbit plane above the external fuel tank was taken as the basic.

One of the key features of MAKS project is the presence of an expendable external fuel tank, and this expands the area of accessible orbital heights. The fractions of the structure mass and payload mass in the total mass orbiter’s injected mass are presented in Table 4. The values of payload mass decrease gradients over the orbital height are displayed for the following three projects: MAKS, AN-225/HOTOL-Interim or MAKS-M and single-stage winged launcher of vertical start (SSTO).

In Fig. 10, dependencies of relative payload mass on the orbit height m_payload/m_payload*, where m_payload* is the payload mass in the reference orbit with the height of 200 km, m_payload is the payload mass at H_orbit > 200 km, are shown. The relations are displayed for three systems indicated in Tab. 4. The line for completely reusable ASS of AN-225/ HOTOL-Interim and MAKS-M type is for brevity designated as HOTOL.

At the indicated in Tab. 4 of fraction of structure mass in the total orbiter’s injected mass, the MAKS orbiter can reach the orbit’s height up to 1,500 km, ASP of HOTOL-Interim type ~ 820 km, and ASP of SSTO type ~ 680 km.

At calculation of economic indices of RSTS of different types, the comparison over a low reference orbit with height of 200 km is aberrant, as the ability of fulfillment of transport operations in working orbits is not reflected. In spite of the fact that SSTO and MAKS orbiter have close payloads in the circle orbit of H_orbit = 200 km, the SSTO already at the height of 680 km shows a zero payload due to its high structure mass fraction (m _structure / m _injected = 85.6%), whereas at the indicated height (680 km) the MAKS orbiter loses only 35% of payload injected into the height of 200 km. It is explained by necessity to boost all the SSTO main fuel tanks’ mass at orbital transitions, whereas in MAKS system the external fuel tank is dropped. As is visible from Fig. 10, at comparing RSTS, it is necessary to take into account not only the payload mass in the reference orbit m_payload*, but also the gradient of payload mass decreasing over the orbital height m_payload/H_orbit.

m _payload / m _payload *, %

H_orbit, km

Figure 10. Influence of orbit height on relative payload mass

The observed now decay of work in researches and developments on advanced space systems abroad is explained first of all by the high costs and high technical risks of creation of such systems as NASP, Sanger, HOTOL and others. As against of these projects, MAKS does not require the solution of such complicated technical problems, as creation of combined propulsion systems, which would combine the air-breathing and LRE modes. To the full extent using the experience of SPIRAL, BOR and BURAN programs, and leading positions of the domestic aircraft industry in the field of heavy transport airplanes, MAKS project looks the most real from all diversity of the researched RSTS concepts. The implementation of this project will cause decrease the cost of flights in space and hazardous affect on the environment, will open the direct access into orbits of low inclinations, will allow to solve efficiently a number of new space missions.

Reference

1. Filat'ev, A.S. The Optimal Space Vehicle Ascent Using Aerodynamic Forces. Cosmic Research, Vol. 29, No. 2. Consultanse Bureau, New York, 1994.

2. Parkinson, R.C., Webb, E. AN-225/HOTOL. AIAA/DGLR 5th International Aerospace Planes and Hypersonic Technologies Conference, Munich, Germany, 30 Nov. - 3 Dec. 1993.

3. Lozino-Lozinsky, G.E., Dudar, E.N., Joyner, R. Comparative Analysis Of Reusable Space Transportation Systems. AIAA 31th Joint Propulsion Conference, San Diego, USA, July 10-12, 1995.

4. Freeman, D.C., Stanley, D.O., Camarda, C.J., Lepsch, R.A. Single-Stage-To-Orbit - A Step Closer. 45th Congress of the IAF, Jerusalem, Israel, October 9-14, 1994.

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