From SPIRAL to MAKS

Thermal Design of the Orbital Planes

Voinov L.P.
The concept of the Orbital Planes’ thermal design is offered. The ways of its realization are shown on the examples of the SPIRAL Orbital Plane and partially on the BOR-4 flying model and the BURAN Orbital Spaceship. The predicted data are confirmed by results of the flight tests.

The development of SPIRAL orbital plane was performed by the group of enthusiasts from the design bureau of A.I. Mikoyan under the direction of the Chief Designer G.E. Lozino-Lozinsky. The study was based on the experience received during the creation of the MIG-25-family fighters with a 3000 km per hour maximum flight speed and 24-km altitude ceiling. These two-engine planes had welded construction from relatively cheap and suitable for manufacturing high-tensile stainless steel. The airfoils of airframe, air intakes and channels in prolonged flight conditions are heated up to temperatures about 300°C, the aft part of fuselage (from hot engine) – up to 500°C, afterburner and jet engine nozzle located in the rear part of fuselage – up to 1000°C. To decrease heating of the rear fuselage steel crimped screen of 0.6-mm thickness with double-sided silver cover was used. It helped to decrease radiant heat flux from the hot engine by ten times.

The orbital space plane differs from the usual planes by three stages of flight: injection into orbit, orbital flight and aerodynamic descend from orbit. Each stage has specific features of gas-dynamic and thermal action on the construction of orbital space plane. That’s why thermal design of the orbital plane was conducted in accordance with specific features of these stages. The descend-from-orbit and the orbital flight stages require special attention. The injection stage must be taken into account when heat protection mass is calculated.

The Thermal Design for the Descent Stage

The principal construction scheme of high-speed plane is mainly determined by maximal allowable construction temperatures and by features of their distribution.

To decrease maximum temperatures to allowable levels and to decrease their gradients on the hottest aerodynamic surfaces a large research and theoretical study was conducted. It was devoted to the study of trajectories and gas-dynamic flows of returned apparatus.

As a results of these researches the conception of the orbital plane thermal designing was suggested. It allows to decrease maximal temperatures to allowable values for high-temperature materials and to decrease temperature differences for reducing of the temperature stresses in hot constructions. Main principles of this conception are the following:

• to fulfil descend at the maximum barometric height (at minimal air density) at the angles of attack alpha = 45°…65°. That conforms to the maximum lift coefficient СLmax;
• to profile load-bearing surface with one critical point, with one flow spreading line, with maximum possible blunt radius of nose part and maximum curvature radius of convex load-bearing surface (flat surface as a limit);
• to perform nose blunt in the form of 55…60° spherical segment with an axis in line with incident flow vector;
• to allow to round sharp edge of spherical nose segment with r ≤ 0,1R;
• to produce internal hollow of spherical segment with maximum radiation clearness for re-radiation of heat radiation, i.e. to produce hollow with maximum filling of internal space;
• to exclude intersection of forward shock with a wing and a fuselage;
• to choose wing sweep angles of wing leading edge and wing trailing edge which provide flow and electron draining after shock wave for extreme reduction of wing edge temperature in the exploitation range of angles of attack and which exclude electron formation at flow action on the edge so that temperature increases a little;
• to foresee wing turn around longitudinal axis of the orbital spaceship for efficient functioning of the wings under subsonic and orbital modes;
• to profile lateral surfaces at negative angle relative to the incident flow vector for considerable reduction of convective flow and formation of heat re-radiation from the hottest lower surface to the practically cold back and later surfaces.

Under offered conditions the chosen form of the orbital spaceship at fuselage length L = 8 m and radiation-transparent blunt with R0 = 1,5 m and at α = 55° and 5…7 tons of landing mass has maximum temperature in blunt zone about 1400° with re-radiation (1600° without re-radiation) and in aft part about 1150° without re-radiation. The maximum temperature of lower wing surface with re-radiation onto the upper wing surface does not exceed 600°C (800°C with re-radiation) at moderate airfoil thickness and blunt radius of the leading edge and flat butt-end of the tailing edge. At that, the wing upper surface temperature (with re-radiation) will be approximately 500°C and at the leading and tailing edges temperature will be less than 600°C.

The mentioned temperatures are realized during the gliding descent from orbit at the angle of attack α =45°…65° and roll angle up to a=60°.

The action time of maximum temperatures is:

• maximum tau max = 15 minutes at alpha = 45°;
• minimal tau min = 7 minutes at alpha = 65°.

On the pointed modes the orbiter spaceship is aerodynamically stable and controllable, the roll angle γ≤60° is allowed to provide required lateral range of descent and to reduce oscillation of the flight altitude during rapid descent.

The main load-bearing element of the orbital spaceship hot structure is the heat protection screen hardly fixed on the farm construction by means of 30 adjustable during the assembling fasteners from niobium alloy with two degree of freedom on spherical ceramic bearings and by means of 7 dual attachment units in a plane of symmetry with one degree of freedom and cylindrical ceramic bearings. In such fixing scheme Y-axial normal aerodynamic (pressure forces) and inertial forces are received by all attachment units while X-axial tangent aerodynamic and inertial loads – in general, by one triple immovable attachment unit located near the cabin.

The farm construction was designed on the basis of high-speed subsonic planes with piston engines. All aerodynamic surfaces with temperature compensators if necessary, ejectable cabin, swivel blocks for wings turning, 4-leg undercarriage, airbreathing engine and maneuvering engines, tanks, equipment bays and other are fixed on the farm.

The fuselage of the orbital spaceship fuselage is designed as non-pressurized (except cabin, fuel tanks and pressurized equipment bays). The pressure inside it is close to barometric atmospheric pressure. This considerably reduces thermal conductivity and, as a result, mass of thermal insulation made of ultra-thin silica materials with 1100°C operating temperature.

The pressure on the lower carrying surface differs from atmospheric pressure in hundred times on modes of maximum heat fluxes and temperatures: Plower surface/Patmospheric = 400 at M = 26 and Plower surface/Patmospheric = 200 at M = 18. Indicated factor have determined the major difference between thermal conductivities of SPIRAL thermal insulation made of ATM-5 material (at t = 100°C L = 0,01 W/m°K) and BURAN tiles heat shield made of TZMK-10 (at t = 100°C L = 0,034 W/m°K).

The toughness of the orbital spaceship heat protection shielding is not on great demand, thus it can be made of clean silica fiber (98…99% SiO2) on the basis of cheap ultra-thin glass thread while tiles heat protection must withstand maximum allowable tear off stress σtear = 2 kgf/cm2. It means that such protection must be made of ultra-clean expensive amorphous quartz using special welding technology for ultra-thin fiber connection. Besides, heat-protection tile must withstand non-transient aerodynamic loads from oscillating surface shock waves on transonic regimes during injection or descending from orbit as well as from the jets of liquid rocket engines during launching and injection.

Due to the large relative pressure difference on the heat protection screen (Plower surface/Patmospheric = 200…400) and for the prevention of snake flows of high-temperature dissociated air the screen must be made in accordance with special technology providing full air-tightness of the lower surface including nose blunt. This air-tightness requirement for the orbital spaceship hot construction is not new and has been already performed during production of afterburners for airbreathing engines, air channels (Рafterburners/Рatmospheric ~ Рchannels/Рatmospheric = 20) and aft parts of hot steel constructions for the high-speed planes. So this requirement is quite realizable.

Aerodynamically smooth shielding (0,7-mm thickness) of the heat protection screen is reinforced by stringer set. The smooth nose blunt is also reinforced by ‘opened’ L-type profiles covered by cases from rhodium foil. The ‘openness’ of L-type profiles guarantees radiation clarity of the hollow nose blunt and internal re-radiation of heat flow from hot blunt surfaces onto cold surfaces. Rhodium cases reduce temperature difference on stringer under re-radiation. The limitation of profile height (h ≤ 20 mm) allows to reduce temperature difference to +-30°C during heating and cooling in the process of descending from orbit.

Non stationary temperature of the heat protection screen is practically the same as the temperature of the thermally insulated surface at quasi steady regime and slightly differs from shielding temperature (within the limits of 2…3%). The heat protection screen, the nose blunt, their stringer sets as well as farm inside hollow blunt are made of niobium alloy. Black coating with emissivity factor ε = 0,8…0,9 is applied on the nose blunt internal surface and on the farm. The farm rod parts next to the screen attachment and hanging units do not heated more then 400 mm along their lengths during the whole flight. They are made of heat-resistant nickel-cobalt alloy with allowable temperature up to 800°C without toughness loss. Other farm rods have maximum exploitation temperature not more than 100…250°C and are made of high-strength low temperature alloyed steel used before in subsonic planes.

To prevent almost completely heat flux penetration from the hot screen onto cold airframe construction, equipment bays, fuel tanks, cabin, farm and other elements, 2-mm thickness high-temperature paperboard with 1500 kg/m3 density and maximum permissible temperature up to 1400…1500°C is fixed on the screen stringers. 30-mm thickness soft fibrous siliceous heat insulation of 160 kg/m3 density is placed on the paperboard. It is pressed by 0,8-mm thickness beryllium sheet. From the outside of the bay this beryllium sheet has silver cover of 20-micron thickness and emissivity ε ≤ 0,03. The temperature of the sheet at the landing doesn’t exceed 400°C thus providing high efficiency of suggested heat protection. Without beryllium sheet the package thickness is approximately 45 mm, in other words the heat protection shielding is considerably increasing in weight. That’s why the exchange of beryllium sheet onto powder type aluminum (sheet of 1-mm thickness) with maximum permissible temperature up to 300°C is foreseen.

The total thickness of heat protection is approximately constant along the whole lower surface and is about 50 mm. The screen hinge hanging allows its temperature expansion (about 40 mm) when descending from orbit.

The upper surface of the orbital spaceship behind a cabin has a temperature that doesn’t exceed 250°C and can be made of titanium or aluminum alloys (by powder metallurgy methods). The inside of the orbital spaceship upper surface has 20-micron silver cover with emissivity ε ≤ 0,03. It practically prevents a heat flux penetration to inside of the plane. It follows that all equipment, fuel tanks, payloads and other orbital spaceship units including pilot cabin are not subjects to the temperature influences that are so opposite for aviation standards. That’s why all complete products made for high-altitude and high-speed planes can be used without any revisions.

The adiabatic wall temperature near the forward cabin illuminators in attachment points of separated flow doesn’t exceed 500°С. The maximum temperature of thick forward glass taking into account transient heating and thermal capacity is about 250°С. This fact allows to use high-temperature glazing made for high-speed planed as illuminators of the orbital spaceship.

The following facts are worth to be stressed because they determine the perfection of the conception for orbital spaceship heat protection.
1. The upper and side surfaces do not require heat protection or heat insulation. It is guaranteed by minimal value of heat flux in stalling zone at maximum hypersonic flow turning when passing obtuse angle. The total area of cold upper and side surfaces is always bigger than for hot surfaces.
2. The hottest lower surfaces of nose blunt and wings have the temperature than is close to the theoretical minimum guaranteed by optimal shape, re-radiation and flight with maximum lift coefficient Суmax.
3. The temperature of the heat protection screen cover is also reduced to minimum theoretical value. This value is assured by one spread line (in a plane of symmetry) with flow lines of maximum length inside of boundary layer including outgoing from a stagnation point with maximum radius of nose blunt and absence of leading edges modes with flow leaking and falling of shock wave.
4. The heat protection package provides simultaneously insulation of hot airframe construction and equipment bays. Thus, the heat insulation mass reduces considerably. Minimal pressure inside airframe also reduces heat insulation mass.
5. The minimal temperatures of the aerodynamic surfaces and small temperature differences allows to produce them from sheet stamped, rolled and continuously welded metal materials.
Hence, the orbital spaceship heat protection system designed on the basis of suggested conception is close to the optimal and its improvement is not advisable.

The upper surface of the orbital spaceship behind a cabin has a temperature that doesn’t exceed 250°С and can be made of titanium or aluminum alloys (by powder metallurgy methods). The inside of the orbital spaceship upper surface has 20-micron silver cover with emissivity ε ≤ 0,03. It practically prevents a heat flux penetration to inside of the plane. It follows that all equipment, fuel tanks, payloads and other orbital spaceship units including pilot cabin are not subjects to the temperature influences that are so opposite for aviation standards. That’s why all complete products made for high-altitude and high-speed planes can be used without any revisions.

All types of temperature influences on plane have been studied for the orbital flight at the minimum permissible altitude H=130 km. The temperature distribution along plane surface depending on total action of all heat fluxes at the angle of attack a= 0 and 55° and the temperature variation in specific points when moving along orbit at night and day are found. At altitude H=130 km it is preferable to fly with zero angle of attack. At that, the temperature near cabin at side, upper and lower surfaces don’t exceed +60°С, at lower surface of airframe +40°С and at its upper surface +20°С.

Aerodynamically stable flight in a molecular flow can be performed only with completely opened wings ψ=90° at the angle of attack α balanced= - 25° and on the tilt wing ψ = 45° at αbalanced= - 45°.

To balance at zero angle of attack it is necessary to open a little an air brake which close the upper fuselage surface in an aft fuselage area and is an excellent balancing flap when flying at subsonic speeds in atmosphere. It considerably improves takeoff and landing characteristics of the orbital spaceship, increases maximum СLmax and eases elevon functioning at СLmax.

The neutral angle of subsonic air brake (balancing flap) is chosen so that to get maximum lift-to-drag ratio (CL/CD)max on subsonic mode. The relative position of tail edges of the balancing flap and the heat protection screen allows to combine hypersonic and subsonic focuses. This problem as we know is still in development. But such combination has been already achieved in SPIRAL project and has been confirmed in subsonic flights of analogue and hypersonic flights of BOR-4 flying models which is geometrically similar SPIRAL model in 1:2 scale.

The subsonic analogue of the orbital spaceship made several flights that confirmed announced takeoff and landing characteristics of the SPIRAL orbiter spaceship. Several falloffs from TU-95 plane were made during flight and independent flights of the orbital spaceship analogue were conducted from one runway to another.

Four BOR-4 orbital flights confirmed all announced SPIRAL hypersonic aerodynamic characteristics including stability and controllability. The external heat flux characteristics were verified in natural conditions of the aerodynamic descending from orbit. For the first time the temperatures on external heat protection surface were measured on the BOR-4 model. Such measuring was made by means of special thermocouple fitted in glass coating of the quartz tiles imitating BURAN heat protection. Relatively uniform temperature was fixed along all the BOR-4 lower surface – 1000…1100°С

Thermal Designing for the Orbital Stage of the Flight

On the orbital stage of the flight the orbital spaceship is under the influence of intensive sun radiation, high-energy molecular air flow (for altitudes 120…140 km), cold space and reflected sun radiation from the Earth surface.

The major method for protection of the orbital spaceship construction against heating and cooling in space conditions are special coatings and specific orientation of the orbital spaceship relative to the Sun and molecular flow vector. The temperature of BOR-4 model on the side and upper surfaces was measured by thermal paints and crystal sensors. The upper surface temperature behind a cabin didn’t exceed 120…250°C and the temperature in attachment area of the secondary flow in a plane of symmetry was less than 300°C, on the side window of cockpit – approximately 400°C and on the front window – less than 600°C

Heat Protection Influence on Aerodynamic Characteristics of the Orbital Spaceship

Very high requirements are established for sub-sonic characteristics of the SPIRAL orbital plane:

• reliable landing on unpaved airdromes with turn-on and turn-off airbreathing engine at landing speed less than 240 km per hour;
• effective and reliable airbreathing engine starting from autorotation at less then three efforts;
• feature to drive to the airdrome using airbreathing engine in case of necessity or if emergency second approach landing takes place;
• effective braking by means of brake flap if overshooting;
• landing technique should not be more complicated than for usual fighters.

Considering that sub-sonic SPIRAL plane was tailless with wings of high sweep and unusual wide load-bearing fuselage, it was obvious that to achieve such landing characteristics were quite complex technical problem. Great contribution in sub-sonic aerodynamic of the orbital plane were achieved thanks to Dr. Samsonov E.A.. He organized aerodynamic models production and their fast testing in wind tunnels.

As a result of sub-sonic studies on the models high efficiency of the balancing flap placed on the upper fuselage surface was proven. It guaranteed maximum lift capacity of the wings at positive deviation angle of elevons (edge downward). It was shown that that balancing flap could be used as a high-performance air brake when it made splitted with perforated upper surface.

The SPIRAL orbital plane was designed for minimal temperatures for flights with maximum lift coefficient and angle of attack α = 55°.

At α = 45°…65° lift capacity changes slightly. That’s why for such angles of attack load-bearing and other surfaces temperatures also changes slightly. However, the further changing of angle of attack relative to the reference value α = 55° ( α < 45° or α > 65°) leads to the considerable increase in surface temperature and overheating of wing leading and trailing edges, side fuselage surfaces and cabin glazing. Increase of angle of slide above calculated value (b = +- 5°) also leads to overheating of side fuselage surfaces and cabin glazing.

Figure 1. Influence of wing flare angle ψ and centering Хc.g. on balancing angle of attack	Figure 2. Dependence between balancing angle of attack and ψ wing flare angle as well as δb.f. balancing flap deviation
S – Orbiter’s wing area in a plane; Swing – wing area with under-fuselage part; L – airframe length

It follows that flights should be conducted when angle of attack α is within the following range 55°…10° and at angle of slide β not succeed …5°. For that purpose great margins in pitch and directional stability should be reserved within indicated values.

Suggested aerodynamic form of the orbital plane has high level of aerodynamic stability.

The angle attack adjustment in SPIRAL at hypersonic mode is performed by wing flare angle alteration without aerodynamic quality losses.

Dependence between balancing attack angle α and wing flare angle ψ as well as gravity center location along longitudinal axis Хc.g. (when Yc.g. = 10%) for descent and orbital flight stages (in molecular flow) are depicted on Figure 1. The upper part of the figure shows how longitudinal centering alteration influences on balancing attack angle

To compare balancing efficiency of the orbital plane depending on wing flare angle and of large area lower balancing flap deviation (36 m2 - 15% of load bearing fuselage surface) a special calculation of balancing, stability margins and lift to drag ratio were conducted. Maximum balancing flap deviation angles were determined during this calculation:

• edge downward delta balancing flap = - 5° - such that not to admit laminar detachment;
• edge upward delta balancing flap = + 15° - limited by engine dimensions.

The comparison shows the advantages of attack angle alteration made by wing flare angle alteration. This method guarantees gains:

• in balancing angle of attack – in two times;
• in longitudinal stability margin – in two times;
• in lift-to-drag losses on balancing Delta(CL/CD)bal=30%.

Figure 3. The wing flare angle and balancing flap deviation influences on the margin of longitudinal stability CmzCL	Figure 4. Dependence between lift-to-drag ratio (CL/CD) values and wing flare angle as well as balancing flap deviation

The balancing angle of attack values αbal at hypersonic descend speeds depending on wing panel flare angle ψ (for the first balancing method see solid line) and on balancing flap deviation angle δb.f. (for the second method see dotted line) are shown on Figure 2. This data shows that wing panel deviation guarantees wide range of αbal values and reference value αbal=55° demands 40° -wing flare angle.

Figure 3 depicts balancing methods influence on CmzCL = ξCmz/ξCL derivative which characterizes margin of static longitudinal stability of the orbital plane.

Figure 4 depicts dependence between lift to drag ratio (CL/CD) values and ψ wing flare angle for the first balancing method (solid line) and δb.f. balancing flap deviation angle for the second method (dotted line). Besides all indicated advantages of deviated wing panel it is worth to mention that constructional design of the balancing flaps is conjugated with splitting of the screen tail-part. This is most heated movable part of construction and necessity of its pressurization in the places of splitting leads to increasing of its mass.

Movable wing is exceptionally effective device for regulation of aerodynamic focus position ХF and longitudinal moment in the focus CmzF. The wing flare angle can also regulate temperature and radio wave transparency of wing leading and trailing edges. It also determines balancing angle of attack and margins of longitudinal stability and changes centering of the orbital plane in wide range.

The SPIRAL project was conducted by high-qualified specialists from design bureau of A.I. Mikoyan. They had wide experience in high-speed planes designing..

SPIRAL design documentation was prepared under the direction of Y.I. Seletsky with his personal participation. Farm fuselage construction and heat protection screen were designed by two chief designers V.F. Pavlov and V.P. Zavgorodsky. Wing hot construction was made by chief designer N.N. Verevkin. The thermal and strength calculations were conducted under the direction of Z.E. Bersudsky. Four-leg ski landing gear (space modification) was designed by chief designer Y.V. Baksht. The design-theoretical documentation for hypersonic aerodynamic and heat exchange was issued by specialists from department headed by L.P. Voinov. He personally participated in orbital plane aerogasdynamic scheme development and performed all experimental-theoretical calculations for injection into orbit, orbital and descending stages. Experimental-theoretical calculations of the orbital plane gas dynamics was done by V.E. Sokolov. He used results of thematic models wind tunnel tests made in TsAGI.

Obtained results were used as an input data for ballistic and aerodynamic calculations and also were utilized by LII for preparing of experimental flights of BOR-2, BOR-3 (1:3 scale) and BOR-4 (1:2 scale) flying models. The flights confirmed all calculations made. V.E. Sokolov conducted calculation of each thermal and force loads affecting the orbital plane. Temperature layout for descending stage was issued by Z.K. Zavyrkina. Transitional construction heating calculations were made by S.F.Teslenko. Temperature scheme at low Earth orbit (H=130 km) if considering molecular flow and sun and Earth radiation was prepared by E.V. Labunskaya. She also calculated polar line and balancing in molecular flow.

Trajectory calculations for injection and descending stages were performed by one of the best design bureau specialist V.A. Shumov. He and Y.I. Seletsky were responsible for SPIRAL balance data during development phase. V.A. Shumov pointed out that reference trajectory for calculation of heat protection thickness and heat analysis would be ‘spring-like’ oscillating on altitude flight trajectory with maximal longitudinal range and flight time without banking (γ = 0 ). He proved that for thermal protection thickness reduction and as a result mass reduction it was necessary to eliminate the “spring” (flight altitude oscillation) using bank angle γ ~ 60° which provides maximum cross range and considerably reduces descending time. It was shown that heat protection mass could be reduced by snake type flight if there was no necessity in large cross range.

The SPIRAL descending trajectory slightly differs from the BURAN trajectory where break impulse directed against velocity vector. SPIRAL break impulse directed at normal to the Earth thus reducing descending time and increases predicted landing accuracy at manual control without use of inertial navigation systems and other devices for trajectory correction.

Several specialists from industrial and academic institutes have participated in project development:
from TsAGI: G.I. Maykopar, V.Y. Borovoy, A.Y. Yashin – in experimental study of heat transfer; G.N. Zamula – in re-radiation inside nose blunt; I.N. Moisheev – in active heat protection system.
from TsIAM: G.G. Tcherny – in numerical calculation of heat exchange.
from NIITP: V.S. Avduevsky, S.S. Tchentsov, N.A. Anfimov – in engineering analysis of heat exchange.
from LII: V.M. Kostylev – in experimental-theoretical analysis of thermal conductivity for low density fibrous materials; L.A. Yumashev – in heat exchange with hydrogen blowing; P.N. Panteleev, A.G. Shibin – in heat exchange for the BOR-2, BOR-3 and BOR-4 flying models; G.P. Vladytchin, A.A. Kondrashev – in dynamics and control development.

The conclusions

1. Heat protection conception for orbital plane was developed. It guarantees minimal temperature of construction if using existing heat protection materials.
2. The conception is based on:

• orbiter flight with maximum lift capacity;
• utilization of ‘lifting body’ configuration with maximum blunt radius of nose and wings which are in ‘draining-from-edge’ mode and not intersected with front shock wave;
• maximum use of heat flux re-radiation from the lower surface onto the upper surface (hollow nose and wings blunt);
• utilization of heat protection screen with inside thermal insulation made of ultra-thin fibers pressed by heat capacity sheet with silver coating;
• utilization of special coatings for heat flux control at orbital flight and descending.

3. To provide stability and controllability and to achieve acceptable temperature and radio transparency of wing leading and trailing edges when descending from the orbit it is advisable to change wing flare angle.
4. The upper surface of the orbiter aft part of fuselage must be equipped with balancing flap which provides stabilization when descending and in sub-sonic mode. It also provides matching of focuses in subsonic and hypersonic modes.
5. Suggested conception of thermal designing was partially used for BURAN and MAKS development. It can be helpful for creation of other space vehicles.