Liquid propellant rocket engine

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises combustion chamber, gas generator, turbopump unit, booster turbopump with gas turbine and heat exchanger. In compliance with this invention, heat exchanges cold inlet is communicated with oxidizer pump outlet. heat exchanges cold outlet is connected via booster turbopump gas feed line with turbopump turbine inlet. Turbine outlet is connected with oxidizer inlet line.

EFFECT: ruled out high-temperature power gas as booster turbopump turbine drive working fluid.

1 dwg

 

The invention relates to rocket engine and may be used in the design of liquid rocket engines (LRE). One of the main challenges when creating LRE is to ensure maximum energy performance. One way to ensure this problem is to increase the differential on the turbine booster turbopump unit (BTA) operating at cryogenic components (e.g. oxygen). BTNA used in LRE as a subsidiary, to provide the necessary pressure at the inlet of the main pump LRE.

Known rocket engine RD-170, which has the power turbines of BTNA oxidant is an oxidizing gas generator taken after the main turbine of the turbopump Assembly (TNA), which is discharged into the path of liquid oxygen and condenses in it (the Path in rocket technology. M: mechanical engineering/mechanical Flight, 2004, p. 120).

The disadvantage of the scheme of this engine is the possibility of condensation of water vapor contained in the gas generator. When condensation of water vapor ice crystals are formed, which clog the filters installed after tha, which can lead to emergency situations in the operation of the engine and even catch fire.

The aim of the invention is to eliminate this drawback, namely the exception I�rnogo high temperature gas as the working fluid of the turbine drive BTNA.

This object is achieved in that in a liquid rocket engine having a combustion chamber, gas generator, turbopump Assembly, booster turbo-pump unit with a gas turbine and a heat exchanger according to the invention the entrance to the cold circuit of the heat exchanger is communicated with the exit of the oxidizer pump and the outlet of the cold circuit of the heat exchanger through a gas supply line to the turbine booster turbopump Assembly with a turbine inlet booster turbopump unit, the output of which is communicated with the inlet line of the oxidizing agent.

The present invention is illustrated in the diagram of the motor shown in Fig. 1, which presents the following units:

1. Input highway fuel.

2. Pump fuel.

3. The fuel supply line.

4. The combustor.

5. The feed line of the fuel in the gasifier.

6. The gas generator.

7. The input line of the oxidizing agent.

8. Pump BTNA oxidizer.

9. The oxidizer pump.

10.Turbine turbopump.

11. The inlet raw gas into the combustion chamber.

12. The exchanger.

13. The inlet of the gas turbine in BTNA.

14. Turbine BTNA.

According to the scheme shown in Fig.1, LRE consists of the input line 1 fuel communicated from the fuel pump 2, which, in turn, informed by the magician�the fuel supply line 3 to the entrance to the cooling jacket of the combustion chamber 4 and the inlet of fuel into the gas generator 5 to the entrance of the gas generator 6. The input line of the oxidizer 7 communicated with the pump inlet, BTN oxidizer 8, and the output from the pump BTNA oxidizer 8 - the entrance to the oxidizer pump 9. In turn, the exit of the oxidizer pump 9 communicates with the entrance to the gasifier 6. The output of the gas generator 6 are communicated with the entrance to the turbine turbopump 10, and the output of the turbine 10 are communicated through the supply line gas generator in the combustion chamber 11 with the chamber 4, wherein the inlet raw gas into the combustion chamber 11 passes through the hot circuit of the heat exchanger 12. The entrance to the cold circuit of the heat exchanger 12 is communicated with the exit of the oxidizer pump 9, and the output from the cold loop of a heat exchanger 12 through the gas supply line to the turbine, BTNA 13 - the entrance to the turbine BTA 14, the output of which is communicated with the input line of the oxidizer 7.

LRE according to the invention operates as follows. Fuel on the input line 1 is supplied to the fuel pump 2, where a portion of it along the fuel supply line 3 is supplied through the cooling jacket of the combustion chamber 4. The other part of the fuel supply line to the gas generator 5 is supplied to the gasifier 6.

The oxidant on the input line 7 is supplied to the pump BTNA oxidizer 8 and then into the oxidizer pump 9, where a large portion of the oxidant fed to the gasifier 6. In the gasifier 6 is produced �teratory gas, which enters the turbine turbopump 10 and leads her into the rotation. Producer gas after the turbine 10 is passed through a hot circuit of the heat exchanger 12, transferring thermal energy component in the cold circuit, is fed into the combustion chamber 4, where it is fired with fuel. A portion of the liquid oxygen after the oxidizer pump 9 is supplied in the cold circuit of the heat exchanger 12, where it is gasified and then supplied to the turbine of BTA 14, causing the latter to rotate. Next, oxygen gas from the turbine BTNA 14 is mixed with liquid oxygen from the input oxidizer line 7, the result is a homogeneous liquid, which is input to the pump BTA 8. The flow ratio of gaseous and liquid oxygen is selected in such a way as to ensure that the process of condensation of gas at the pump inlet without effervescence of the main flow of fluid, i.e. to satisfy the condition of excess pressure in the mixture above the saturated vapor pressure.

Thus, the negative pressure at the outlet of the turbine BTA 14 and as a consequence higher differential on it directly without the use of producer gas of high temperature. This solution eliminates the presence of gaseous oxygen coming from the turbine BTA 14, water vapor and other components, which can condense into solids.

Use�the Oia of the present invention will improve the internal energy of the rocket engine and thereby improve its operational performance (resource, efficiency and reliability).

Liquid propellant rocket engine having a combustion chamber, gas generator, turbopump Assembly, booster turbo-pump unit with a gas turbine and a heat exchanger, characterized in that the entrance to the cold circuit of the heat exchanger is communicated with the exit of the oxidizer pump and the outlet of the cold circuit of the heat exchanger through a gas supply line to the turbine booster turbopump Assembly with a turbine inlet booster turbopump unit, the output of which is communicated with the inlet line of the oxidant.



 

Same patents:

Turbo-pump unit // 2548331

FIELD: engines and pumps.

SUBSTANCE: turbo-pump unit (TPU), incorporating a rotor and a stator, according to the invention, is fitted with the controlled plunger with a working end face placed in the stator and mobile in the axial direction, and on the rotor the companion end face is provided, and in the working position of the plunger both end faces are in power contact for keeping of the rotor in motionless position. Besides, the stator is interconnected with the plunger by a tight bellow valve interconnected with the control pressure nipple, from the side, opposite to the working end face.

EFFECT: prevention of rotation of TPU shaft in the autorotation mode at purges and technological works on the engine for increase of resource of operation of bearings and other TPU elements.

2 cl, 1 dwg

FIELD: engines and pumps.

SUBSTANCE: rocket engine contains a combustion chamber where borane, or silane, or phosphene, or germane, or other hydrides with a positive enthalpy of formation from simple substances or their mix are fed. The named above substances a fed at the temperature ensuring self-sustaining course of reaction of their thermal decomposition due to heat of exothermic reaction. Another invention of the group relates to the liquid or solid fuel rocket engine into the combustion chamber of which in addition to stoichiometric composition of the main fuel borane, or silane, or phosphene, or germane, or other hydrides, or methane are fed. Another one invention of the group relates to the solid fuel rocket engine in which solid hydrides in addition to stoichiometric composition of the main fuel are a part of solid rocket fuel.

EFFECT: group of inventions allows to increase a specific impulse of a rocket engine.

9 cl

FIELD: engines and pumps.

SUBSTANCE: this process proceeds from addition of polymer anti-turbulence additive (ATA) as agent reducing hydrodynamic loss in kerosene fuel line. It comprises feed of oxidiser and fuel to engine combustion chamber to get combustion products and their expansion in jet nozzle to create engine thrust. Note here that polymer ATA is added from extra tanks and mixed with pure kerosene flow fed to engine inlet fuel line at engine starting and operation in the mixer arranged in said line. Said polymer ATA represents the solution of polyisobutylene in concentration of 0.6-0.8% of pure kerosene bulk or the solution of high alpha-olefins in kerosene with concentration of 0.6-0.8% of pure kerosene bulk. Rocket engine plant comprises liquid-propellant rocket engine with turbopump fuel component feed system, oxidiser ( liquid oxygen) tank and fuel (pure kerosene) tank and fuel lines to connect said tanks with engine. Besides, it comprises extra tank filled with the solution of aforesaid polymers and displacement system for feed of said polymer in fuel inlet line for its mixing with pure kerosene during engine operation. Said extra tank has diaphragm to divide its in two chambers so that fluid chamber is communicated with the engine inlet line. Second, gas chamber is communicated with high-pressure cylinder via the valve and pressure control valve.

EFFECT: higher payload to be placed in orbit.

5 cl, 3 dwg

FIELD: motors and pumps.

SUBSTANCE: power pack comprising the tanks with circuits 1, 2, boosting system 3, gas-producer 4 with ignition device 5 for transforming of liquid cryogenic oxidant into gaseous oxidant with preset temperature and storage receiver 6 for gaseous oxidants as a fuel for motor units 7, contains the heat exchanger 8 for transforming of liquid cryogenic fuel into gaseous fuel with heating up to preset temperature comprised by heat transfer path into the circuit downstream the gas-producer 4, by the heat receiving path - into the circuit for supply of cryogenic liquid fuel, storage receiver 9 for gaseous combustible fuel for power supply of motor units 7, comprised in the circuit downstream the heat receiving path of the heat exchanger 8, gas liquid mixer 10, comprised in the circuit between the output of the heat receiving path of the heat exchanger 8 and input of the storage receiver 6 for gaseous oxidants, meanwhile the fluid input of the mixer 10 is connected with the circuit of liquid oxidant supply to the gas-producer through the pipeline 11 with the adjusting throttle washer 12 installed.

EFFECT: improvement of reliability of power packs of reaction control systems using liquid cryogenic fuels.

1 dwg

Propelling device // 2532326

FIELD: engines and pumps.

SUBSTANCE: propelling device comprises a body, a cone-shaped combustion chamber, an exhaust pipe, two spring valves between the exhaust pipe and the combustion chamber, a control unit with hydraulic outlets.

EFFECT: invention makes it possible to increase reliability of operation of a jet engine without speed reduction.

1 dwg

FIELD: engines and pumps.

SUBSTANCE: liquid propellant rocket engine comprising a combustion chamber, a turbopump set, a drainage cavity connected with a drainage pipeline, at the same time the drainage cavity is located between an oxidant pump and a turbine, and the drainage pipeline is equipped with a gas ejector, at the same time the gas ejector is connected by a pipeline with a cavity downstream the turbine. The pipeline comprises a valve and a throttle.

EFFECT: increased efficiency of a system of LPRE cavities draining and removal of fuel components from them.

2 cl, 2 dwg

FIELD: chemistry.

SUBSTANCE: described is a method of increasing energy characteristics of liquid rocket engine, working on fuel components being liquid oxygen and hydrocarbon combustible, with kerosene with a liquid additive, representing a solution of highly molecular polyisobutylene (PIB) with a medium-viscosity molecular weight from 3.1·106 to 4.9·106 in kerosene in an amount, providing concentration of polyisobutylene in kerosene from 0.015% to 0.095% of kerosene weight, being applied as the hydrocarbon combustible; cutting of an impeller of a combustible pump of the engine turbopump unit is performed, with an external diameter of the impeller D2 being determined by formula D2=D1(1BC1+A)0,5, D1 is an external diameter of a working wheel of a standard combustible pump; A is a relative increase of the combustible pump pressure in operation with the PIB; B is a relative decrease of hydroresistance of a tract of a chamber regenerative cooling caused by the PIB impact; C=ΔPcoolΔPp is a ratio of hydroresistance of the tract of the chamber regenerative cooling to pressure of the pump for the supply of the component without the PIB, in order for the value of a weight ratio of components (Km) when the engine operates in nominal and forced modes with an application of kerosene with the liquid additive PIB to remain equal to the value Km when the engine works on pure kerosene.

EFFECT: increase of energy characteristics of the LRE.

2 dwg, 3 tbl

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises liquid or solid propellant fuel wherein oxidiser and/or combustible includes fixed nitrogen as well as fine or fixed boron. Note here that boron atoms-to-nitrogen atoms ratio makes 1:1 with departure of ±20%. Rocket fuel has fuel excess relative to oxidiser.

EFFECT: higher fuel heat generation.

9 cl

FIELD: engines and pumps.

SUBSTANCE: invention relates to liquid-propellant rockets, accelerating units and can be used at starting the engines when liquid-propellant store residues do not exceed 3% of initial value. Proposed method consists in gasification of liquid residues of unusable fuel reserve in oxidiser and combustible tanks, generation of braking pulse by their combustion in combustion chamber of gas rocker engine and high-rate blowdown of combustion products into space. In compliance with this invention, solid-propellant gas generating compounds (SPGGC) are used for gasification of unusable rocket propellant reserve. SPGGC is fed to oxidiser tank with excess oxygen while SPGGC with limited content of oxygen is fed to fuel tank. Note here that chemical composition and mount of SPGGC at minimum possible residues of unusable rocket propellant components are defined proceeding from preset characteristic speed: ΔVΣpreset=ΔVresSPGGC+ΔVSPGGCSPGGC, where ΔVΣpreset is characteristic speed, ΔVresSPGGC is pulse developed owing to minimum unusable residues of rocket propellant in both tanks required for their oxidation, ΔVSPGGCSPGGC is the pulse developed only by combustion of SPGGC gases in gas rocket engine. Proposed device comprises engine with oxidiser and fuel tanks, tank supercharge system, has rocket engine with power supply system and rocket propellant component residues gasification system. Note here that engine plant is equipped with solid-propellant gas generators with their outlets connected with gas feed devices. Said gas generators are equipped with pyro membranes fitted in fuel tanks with residues of liquid rocket fuel.

EFFECT: higher efficiency.

3 cl, 1 dwg, 1 tbl

FIELD: engines and pumps.

SUBSTANCE: proposed method proceeds from gasification of rocket propellant components (RPC) to be fed into combustion chamber. Note here that after outage of liquid-propellant mid-flight engine RPC gasification system is actuated. Supercharge gas is fed into balloons with extra RPC. Redox gas generators are used to feed heat carriers into tanks with RPC residues depending upon specific fuel kept in tanks.

EFFECT: higher power efficiency and environmental safety, better operating performances.

1 dwg, 1 tbl

FIELD: rocket technology; heating gases using heat produced in nuclear fusion.

SUBSTANCE: proposed method is characterized in that gas is introduced in at least one chamber. The latter has wall coated with disintegrating material. This material is exposed to neutron flux to induce disintegration into fragments within chamber. Mentioned wall is cooled down on rear end relative to chamber and mentioned coating. In addition, device implementing this method is proposed. Gas heating device has at least one gas holding chamber. It has wall coated with disintegrating material and facility for exposing disintegrating material to neutron flux so as to induce and emit disintegration fragments within chamber. Device is designed to cool down mentioned wall on rear end of chamber and mentioned coating of disintegrating material. In addition, space engine using mentioned method for gas heating is proposed. This space engine has gas heating device and facility for exhausting hot gas into space to afford thrusting. Alternative way is proposed for gas heating by using nuclear fusion reaction suited to space engines for thrusting.

EFFECT: facilitated procedure of gas heating.

42 cl, 24 dwg

FIELD: rocketry and space engineering; rocket pod engine plants.

SUBSTANCE: proposed engine plant includes propeller tanks (oxidizer tank and fuel tank), cruise engine, actuating members and high-pressure gas bottles. Oxidizer and fuel tanks are filled with low-boiling and high-boiling components, respectively. High-pressure gas bottles are installed in oxidizer tank. Rocket pod engine plant is provided with pipe lines mounted on fuel tank by means of brackets forming heat exchange unit. Pipe line inlets are communicated with outlets of high-pressure gas bottles and their outlets are communicated with actuating members of engine plant.

EFFECT: reduced mass and volume of high-pressure gas bottles and consequently reduced mass of rocket pod.

1 dwg

FIELD: rocket-space equipment, mainly means and methods for water supply to low-orbital spacecraft.

SUBSTANCE: the offered method provides for use of the energy of formation of the raw material, in particular, of water from the fuel components for increasing the efficiency of the means of its injection into orbit. The offered rocket power plant has a chemical reactor, in which the given product is formed, as well as a heat-exchange unit, in which the heat of the chemical reaction is transferred to the fuel components. The latter results in the growth of the power plant specific impulse. The reaction product is cooled, and a condensate (water) is obtained which is accumulated in the storage tank. The offered rocket may use one of the cleared fuel tanks for accumulation of condensate. The offered transportation system includes the offered rocket, orbital station equipped with a system of water processing to fuel components, and means of delivery of the space vehicle to the station together with the non-filled boosting unit. The offered transportation-fueling station includes also an orbital fueling complex. Space vehicles injected into high-altitude orbits, in particular, into a geostationary orbit, as well as space vehicles returning on the Earth, may be refueled there. At injection of the space vehicle into a geostationary orbit the dependence of the efficiency of injection on the latitude of the cosmodrome is essentially reduced (by 2-3 times).

EFFECT: reduced cost of supply of the orbital stations and cost of injection of the space vehicle into a geostationary orbit, as well as into other trajectories, reduced dependence of the cost of injection of the space vehicle into a geostationary orbit on the latitude of the cosmodrome.

19 cl, 3 dwg

FIELD: aircraft industry; rocketry.

SUBSTANCE: invention relates to design of liquid-propellant rocket engines. Proposed liquid-propellant rocket engine without afterburning of generator gas contains regenerative cooling chamber 1, turbopump set 2 with gas generator 3 to drive turbine 4, two flow rate controls and two nozzles 9, 10 installed in pressure main lines 11, 12 of pumps of turbopump set 2. Sensing elements of spools 5, 6 of controls communicate through pipelines with inputs of nozzles 9, 10 and their minimum sections. According to invention servo-actuate restrictor 14 of control, playing the part of thrust control, installed in feed main line 12 of one of propellant components into gas generator 3. Restricting element of servo-actuated restrictor 14 communicates through pipeline 21 with pressure main line 12 of pump of said component after nozzle 10, and pipeline 22 delivering second component into gas generator 3 is connected with pressure main line 11 of pump of said component after servo-actuated restrictor 13 of control playing the part of propellant components flow rate ratio control.

EFFECT: improved energy-mass ratios of engine, provision of constant propellant components flow rate through engine and thrust irrespective of ratio of components passing through engine.

1 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocket-propelled vehicles, particularly, to the gas duct of liquid-propellant rocket engines with after-burning. The aforesaid gas duct comprises the outlet manifold of the main turbo-pump unit, a bent pipeline and a swinging assembly. The aforesaid bent pipeline is coupled with the outlet manifold and the said swinging assembly is linked with the engine chamber. Note here that the aforesaid swinging assembly is furnished with a bi-degree universal joint and the joint of the swinging assembly with the engine chamber and bent pipeline represents a flange coupling incorporating a metal T-shape gasket furnished with a load-bearing ring with two flexible springs provided with mountain-like ledges. Note also that the aforesaid one-piece bent pipeline is made from a heat-resistant nickel-alloy, while the bent pipeline flange represents a load-bearing belt with a developed end face surface for the engine frame support to be attached thereto. The aforesaid T-shape taper gasket springs feature the thickness varying over their length, while their length L-to-mean thickness δ ratio makes L/δ ˜8 to 10 and the angle α of the spring taper surface inclination to the flange coupling axis makes 1.5 to 2.5 degrees. The flexible spring OD including the aforesaid mountain-like ledges exceeds the ID of the flange coupling sealing surfaces by 0.1 to 0.2 mm. All parts of the gas duct are made from the EK-61 heat-resistant nickel alloy. The propose invention allows a higher tightness of the fixed joints and pipelines carrying high-temperature high-pressure oxidising medium.

EFFECT: improved performances due to ease of uncoupling gas duct from engine chamber and bent pipeline.

7 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: in method for compensation of differences in physical properties of fuel components based on matching of operation modes of universal liquid-propellant rocket engine supply units, according to invention for generator-free engine with separate turbine pump (TP) during its transfer from hydrogen to liquefied natural gas (LNG) (methane), at first fuel (LNG, methane) flow is increased to required value for provision of reliable cooling of chamber, after cooling prior to fuel supply to turbine of TP its total flow is divided into two parts, one of which is supplied to TP turbine, and the other one is discharged, at that after TP passing, fuel fission process is repeated, at that its one part is sent for combustion in combustion chamber, and the other is discharged or sent for further use. Discharged parts of fuel flow may be used as working fluid, for instance, for steering nozzles, for turbine of engine swinging system, for supercharging of tanks, repeatedly as working fluid of chamber fuel and/or propellant pump. Invention provides for operation of engine both on fuel components "oxygen+hydrogen" and also on fuel "oxygen+liquefied natural gas" (methane).

EFFECT: reduced cost of engine and expanded field of its application.

7 cl, 4 dwg

Rocket engine unit // 2381378

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry and can be used in designing rocker carrier first stages with multi-tank propellant compartments with wrap-around arrangement. Engine unit comprises multi-tank propellant compartment and fluid propellant rocket engines, every engine being communicated, via feed lines, with adjoining tanks. One of the engines communicates, via feed lines and booster pump units, with all tanks.

EFFECT: synchronised utilisation of propellant components from like tanks without introducing disturbing torques to rocket.

1 cl, 1 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry, particularly to liquid propellant rocket engines operated on three fuel components, i.e. cryogenic oxidiser, hydrocarbon fuel and liquid hydrogen. Proposed engine comprises at least one combustion chamber with jet nozzle, regenerative cooling system, gas generator, and turbopump unit comprising turbine, oxidiser pump and fuel pumps. It differs from known designs in that said turbopump unit comprises two fuels pumps and two extra fuel pumps designed to operate on first fuel and second fuel. Note here that second fuel pump and additional second fuel pump are arranged below oxidiser pump. Downstream of fuel pumps, first and second fuel valves are arranged connected, via electric line, with synchronisation device. Proposed engine incorporates also control unit connected with aforesaid synchronisation device. Method of operation of above described engine comprises feeding fuel and oxidiser into gas generator and combustion chamber, igniting them and exhausting combustion products via jet nozzle. In compliance with this invention, first fuel utilised, second fuel is fed into gas generator and combustion chamber. Prior to feeding second fuel, fuel pipelines and nozzle regenerative cooling systems are blown down to remove first fuel residues.

EFFECT: improved operating performances of liquid propellant engine in wide range of flight conditions at various altitudes.

5 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry, particularly to liquid propellant rocket engines operated on three fuel components, i.e. cryogenic oxidiser, hydrocarbon fuel and liquid hydrogen. Proposed rocket comprises first- and second-stage rocket units connected in parallel, oxidiser and fuel tanks coupled by power assemblies and equipped with at least one first-stage engine and one second-stage engine. In compliance with this invention, second-stage unit comprises second fuel tank, every second-stage engine incorporates combustion chamber and fuel feed turbopump unit. Proposed engine comprises at least one combustion chamber with jet nozzle, regenerative cooling system, gas generator, and two turbopump units comprising turbine, oxidiser pump and fuel pumps. In compliance with this invention, outlets of all pumps communicate, via gas duct, with gas generator outlet communicates with every combustion chamber. Method of operation of above described engine comprises feeding fuel and oxidiser into gas generator and combustion chamber, igniting them and exhausting combustion products via jet nozzle. In compliance with this invention, first fuel utilised, second fuel is fed into gas generator and combustion chamber. Prior to feeding second fuel, fuel pipelines and nozzle regenerative cooling systems are blown down to remove first fuel residues.

EFFECT: higher thrust-to-weight ratio, improved operating performances.

12 cl, 8 dwg

FIELD: engines and pumps.

SUBSTANCE: invention can be used during development of liquid-propellant engines (LPI) for carrier rockets (CR). Method consists in the fact that acceleration pulse is created owing to combustion of fuel components in ignition device (ID) and supply of its combustion products to chamber nozzle. Ignition device is tripped after the required pulse is obtained by the mixture in CR tank. At that, combustion products are supplied to the combustion chamber nozzle together with their ballasting, e.g. with fuel which is first passed through the cooling path of the chamber. The proposed method is implemented in LPI containing combustion chamber with ID, nozzle, turbo-pump unit, automation and control units, which, according to the invention, is equipped with an additional line with the valve for ballasting of ID combustion products, which connects the outlet of the cooling path of combustion chamber to its mixing head.

EFFECT: simplifying the design and reducing power consumption.

3 cl, 1 dwg

Up!