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In NPO MOLNIYA, the alternative versions of two-stage launch systems of vertical start, in particular, the system with a reusable winged booster of the 1-st stage on the basis of BURAN Orbiter design and with expendable rocket block as the 2-nd stage booster were also studied.

Single-stage ASP of vertical start and horizontal landing (ASP-O) was studied together with NPO ENERGOMASH. The similar project (SSTO) is accepted now as the base in the American program of RLV reusable launch vehicle, which provides the study of three concepts:

• conical SSTO of Delta-Clipper with vertical rocket landing;

• winged SSTO of vertical start and horizontal landing;

• SSTO of ‘lifting body’ scheme of vertical start and horizontal landing.

As is known, the SSTO ‘lifting body’ scheme project, designed by Lockheed Corporation and providing advanced creation of a diminished experimental FV-demonstrator X-33, won in the competition.

The optimal performances of ASP-O single-stage launcher (Fig. 6) are achieved at using three-component LRE.

Figure 6. ASP-O project (SSTO)

It is necessary to note that three-component engines provide greater advantage in the ASP-O and SSTO projects (as compared to dual-component Space Shuttle SSME), than in two-stage systems. At fixed ASP take-off mass of 550 tons, dry mass of the launcher is reduced by 18%, which results in increase of payload mass from about two to ten tons. The very close results of three-component LRE for SSTO efficiency assessment were obtained in the USA [4].

The launch characteristics of the winged single-stage launcher can be improved at the expense of using the aerodynamic lift capacity at implementation, after the vertical take-off, of more flat trajectory with smaller fuel expenditures. In Tab. 3, the main mass characteristics and components of losses in characteristic velocity are shown at SSTO launch under the scheme, conventional for the vertically starting launcher with gravitational turn and with optimal usage of aerodynamic forces. At take-off mass of SSTO equal to 550 tons, the flat launch due to the essential decrease of gravitational losses allows to increase payload mass by more than 10%.

In Tab. 3 the data for horizontally launched systems (MAKS and AN-225/HOTOL) are shown for comparison.

As compared to MAKS second stage, HOTOL-Interim has the higher lift-to-drag ratio that decreases the requirements to intensity of CP pre-launch maneuver and provides more flat injection trajectory with smaller losses of characteristic velocity. It results in increase of injected mass into a low-height orbit. But on payload mass the completely reusable ASS loses as compared to MAKS system because of the greater fraction of the structural mass in injected mass, and the loss in payload mass increases with the increase of the orbital height. The results in Tab. 3 correspond to the reference orbit with the height of 200 km and inclination of 51^o.

Table 3 displays the energetic contribution of the carrier-plane as the first stage of launching system. This contribution to launch is not explained only by the initial speed (of approximately 190 m/s). Gravitational losses decrease results in the ideal speed gain of 560…650 m/s as compared to the conventional vertical start, and the gain is higher at the increase of the lift-to-drag ratio of the 2-nd stage of ASS.

Single-stage winged launcher with LRE, started horizontally from the airfield accelerating trolley (ASP-R) has appeared the least attractive of all studied versions. Essentially, such RSTS (Fig. 7) is a two-stage design since the accelerating unit actually represents the ground-based booster of the 1-st stage. The disadvantage of such start consists in additional energy expenditures connected with ASP injection into the plane of launch, when the direction of runway does not coincide with the launch azimuth. Naturally, the start from the mobile ground-based platform decreases the undercarriage mass. But ASP wing in this case is calculated under the conditions of horizontal start from the ground, as against SSTO and MAKS and AN-225/HOTOL-Interim systems, in which the size of a wing of a reusable orbital stage is determined according to the landing conditions.

Single-stage ASP-R launched horizontally from the mobile platform can represent a practical interest, if this platform is not a ground-based ‘cart’ or ‘trolley’, but AN-225 airplane. But in this case ASP-R is transformed into a system of ASP-D type.

At preparation of the technical proposal on reusable winged ASP, the horizontal type of start was studied at different values of Mach number (М₀) and flight height (Н₀) at the beginning of autonomous phase of launch of an aerospace plane with LRE:

• start ASP-R from the ground-based accelerating device at М₀ = 0.5 and Н₀ = 0;

• start ASP-D from the subsonic carrier-plane at М₀ = 0.7 and Н₀ = 8…9 km;

• start ASP-H from the hypersonic booster plane at М₀ = 5.0 and Н₀ = 25…30 km.

Figure 7. ASP-R with accelerating unit

The dependence of payload mass m_pl on the radical square from specific mechanical energy E₀ at the start of autonomous phase of launch at fixed launching ASP mass of 250 tons is displayed in Fig. 8.

This practically linear dependence displays that the start from the ground-based accelerating device essentially loses to the air start by power capabilities. The hypersonic air start increases considerably the payload mass as compared to the start from subsonic CP. However, implementation of projects of ASP-H type is the task of more long-term future, and at existing technological base the most promising path is the creation of a partially reusable MAKS system with an expendable ET, which provides as contrasted to ASP-D twice to increase the payload mass injected into a low Earth orbit inside payload bay of the reusable orbital plane.

Comparing now results of researches, corresponding to all the listed projects, we see once again that MAKS system represents the most reasonable compromising proposal for the nearest period. This project implementation could become the basis for future creation of completely reusable launching systems.

But, as any launching system, MAKS has perspectives of improvement of its basic performances in the process of aircraft and rocket-engineering technologies development. Application of new materials and onboard equipment will permit to decrease the dry mass of ET and OP. The characteristics’ essential improvement can be reached in case of creation of a new special heavy HERACLES carrier-plane with weight-lifting ability up to 450 tons, which project is developed at NPO MOLNIYA. This plane (Fig. 9) is design according to ‘tri-plane’ scheme with two fuselages, between which the payload on the external attachment is placed: second stage of ASS, bulky commercial freights, cargo container with freights or removable passenger cabin.

Figure 9. Advanced HERACLES heavy carrier-plane (with cargo container)

In case of HERACLES application, a completely reusable aerospace system of ASP-D type can provide the same payload mass, as the MAKS partially reusable system with AN-225 carrier-plane.

Design-Ballistic Researches
under MAKS Project

One of the first projects of an aerospace system was researched at NPO MOLNIYA in 1977…1978 within the framework of ROSA R&DW. This ASS represented a subsonic carrier-plane and the second stage consisting of a small orbital plane (OP) of ‘lifting body’ scheme, in tandem joint with a rocket booster (RB). The second stage (RBOP = RB + OP) reminded very much the second stage of SPIRAL system. Later on, this scheme was studied within the frames of works on ASS-49 and BIZAN systems, underwent various changes and eventually was transformed into the now widely known MAKS project on the basis of AN-225 heavy carrier-plane. The detailed description of ASS-49 and BIZAN systems is made in article ‘Two-Stage ASS of Horizontal Start’ of this book.

A disadvantage of the scheme with RBOP was the presence alienated fields at the distance of approximately 2000 km from the point of start for drop of expendable rocket booster. In the now known MAKS scheme, the external fuel tank is dropped in the antipodal region of the World Ocean.

The higher scale of MAKS reusability MAKS as compared to ENERGIA-BURAN system is achieved due to the absence of expendable rocket booster at the first stage, instead of which CP is used, and arrangement of the main LRE on the reusable orbital plane itself.

Certainly, the higher economic indices, as contrasted to conventional launchers, is not a sole advantage of ASS with subsonic CP. The available CP ranges, in particular, with using the air refueling, permit to depart to the zone of southern latitudes and to provide start directly in the plane of near-equatorial orbits. This new quality of the launching system is especially valuable for Russia and other countries with a similar geographical position.

The air start mobility provides also the second important advantage - high efficiency of injection into the scheduled orbit and docking with the orbital object.

Absence of alienated fields on the ground and application of ecologically pure components of propellant increase the attractiveness of MAKS concept.

Certainly, the now known form of MAKS has not appeared immediately. The path of search, doubts, errors and finding the optimal solutions was long. The accumulated experience of designing on such complicated system and creation of extensive cooperation between the organizations participating in these works are of great value for the future of domestic aircraft and space-rocket branches.

The major question in formation of MAKS system configuration was selection of LRE for the main propulsion system. Already at the early stages of researches on ASS, it became clear that kerosene LRE do not provide the required level of payload mass, because of low specific impulse. To the beginning of the stage of draft designing, the general configuration of MAKS was formed with three oxygen/ hydrogen NK-45 engines developed, by Samara NPO TRUD with thrust in vacuum of 90 tons each. The take-off mass of the second stage was of 250 tons, maximum payload mass on the orbit with a height of 200 km and inclination of 51^o was 7 tons. The cycle of researches was conducted later, which have shown that the best characteristics of MAKS launch are achieved in case of application of the three-component dual -mode RD-701 engine. The gain as compared to the oxygen/hydrogen engines of NK-45 type in three-component LRE version is achieved, despite of decrease of average over the trajectory specific impulse, as a consequence of essential decrease in the hydrogenous combustible and external fuel tank mass. Besides, being the engine of a new generation, RD-701 has a lower specific mass. Three-component LRE has allowed to reduce ET geometrical sizes and to improve aerodynamics of the MAKS second stage. Together with NPO ENERGOMASH, the optimization was conducted on hydrogen percentage in propellant consumption and on ratio of the thrust levels at the first and second modes of RD-701 operation.

The task of selection the MAKS second stage optimum layout, which consisted of two objects - external fuel tank (ET) and orbital plane (OP or Orbiter), was very difficult. Versions with arrangement of the oxidizing agent in the forward part of ET were considered at the beginning. In this case because second stage of MAKS (OPET = OP + ET) had excessive static stability. For decreasing the OPET aerodynamic longitudinal moment on diving it was necessary to fold completely the declined outer wing consoles of OP, in order that they did not create the lift affixed in a rear part of OPET. However, such solution contradicted to the already known principle of using the aerodynamic lift at the space systems launch. In this case the very high values of trajectory angle (of 40…45^o) were required at the moment of OPET separation from carrier-plane, because of low lift-to-drag ratio of the 2-nd stage. Trajectory inclination at such angles could be provided only due to the prolonged LRE functioning until separation or with the help of additional LRE placed in the tail part of carrier-plane. At that in order not to spend a lot of fuel from ET before separation, it was necessary to install additional tanks with rocket propellant components inside the carrier-airplane. The developers of carrier-plane raise an objection against such solution.

One variant to solve the problem was the installation of lifting aerodynamic surface (small destabilizing wing) at the nose of external tank ‘canard’. This has allowed to decline the outer orbiter’s wing consoles at an angle of approximately 45^o from vertical plane and to indemnify the weight of the forward wing at the expense of increase in the OPET mass injected into an orbit. The forward wing on ET don’t only increase the aerodynamic lift by itself, but also allows using the lift from the orbiter’s outer wing consoles. Such solution had a number of essential disadvantages: stall from the forward wing influenced on streamlining flow on outer orbiter’s wings, the bending moment which acted on the ET construction was increased, ET structural mass and cost increased too.

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