ES2691990A1 - Tangential flow non-positive displacement motor (Machine-translation by Google Translate, not legally binding) - Google Patents

Abstract

The invention constitutes a tangential flow non-positive displacement motor; As a turbomachine, it integrates into fewer elements several stages that the previous technique carried out in several separate elements. The invention is based on a rotor disc, in which several rotating nozzles are integrated, in which in turn are integrated into each, a combustion chamber, fuel injection and control system, as well as oxidant input. At the same time, the turbine integrates the implicit high pressure fuel injection pump to the rotation thereof, the fuel being introduced by the inner body of the turbine through its axis and accelerated by the centrifugal force. (Machine-translation by Google Translate, not legally binding)

Description

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D E S C R I P C I O N

MOTOR OF NON POSITIVE DISPLACEMENT OF TANGENTIAL FLOW

TECHNICAL SECTOR

The present invention is mainly framed within the industry of turbomotor devices, turboprop engines and internal combustion engines applicable in the aerospace industry, as well as part of processes of conversion of fuels into thermal-mechanical energy applicable in many industries, power generation electric, propulsion of land vehicles and internal combustion machinery.

BACKGROUND OF THE INVENTION

The state of the prior art comprises a wide range of turbomachines, mostly implemented in the aeronautical industry. By dividing them into large groups, the current technique includes turbomotors, turbo-reactors and turboprops. All of them are based on the mechanical use of a fluid that undergoes a change in temperature, defined thermodynamically by the Brayton cycle and where its thermodynamic performance of Carnot establishes that, = 1 - -, being 7 \, initial temperature of the fluid and T2 the final temperature of it.

Turbomotors, turbo-reactors and turboprops, are very similar machines between them, whose main difference lies in how they use their mechanical performance. All of them consist of what is known as the gas generator block, which consists of capturing an oxidizer from the atmosphere (oxygen), compressing it together with the nitrogen in the air, mixing it with a fuel, combustion, raising the temperature / volume of the fluid and extract a mechanical performance in one or several blade turbines that at least cover the energy invested in the process, the blade turbine being defined as that device that is capable of converting the kinetic energy of a fluid into rotating mechanical energy. Once the energy deficiencies in the gas generator block are covered, the energy excess is used according to the mentioned types, the turbomotors use more turbine stages to transmit said mechanical energy to a device, the turboprops in the same way to a propeller type helix or fan, and turboreactors to accelerate the fluid in a nozzle with the remaining kinetic energy in order to produce a thrust.

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In this definition, all those processes in which their energy balance is negative, that is, require energy for their development, are named as energy deficit, which is of vital importance to understand the performance of a turbomotor and the advantage of this invention. . In a current turbomotor, the negative energy balance elements are diverse, but mainly the compression phase, which consumes approximately two thirds of the mechanical energy generated in the gas turbine, a figure that does not match the energy needed to compress mass air needs at stoichiometric levels of the fuel, and which has its explanation that the turbomachines in the present technique compress around 40% more air than is necessary for the complete combustion of the fuel, the reason is that the turbine blades of gas would not withstand combustion temperatures at exclusively stoichiometric volumes, so this excess air is used among others to lower the temperatures of the blades to acceptable values, both by lowering the temperature of the primary combustion flow, as flow secondary to cool the blades directly, which in the end results in a series of losses both mechanic as well as thermodynamics, or what is the same, a notable reduction in efficiency, in addition to a constructive set of volume and greater weight.

On the other hand there are rocket engines, although they are very different internal combustion machines than those mentioned above, as a thermal machine they have the same phases, compress an oxidant and a fuel or use a solid or liquid monopropellant, which when reacted in a combustion chamber, mechanical energy from that temperature variation is extracted. Although, in this case, turbines are not used for their extraction of mechanical energy, but devices called convergent-divergent nozzles separated by a narrow point called the throat, which are capable of converting the thermodynamic expansion of the fluid resulting from combustion into a mechanical acceleration in the form of useful propulsion work. Although, these devices under flow rates in supersonico-hipersonico regime are some of the most efficient thermodynamic machines and that are closer to the maximum Carnot performance, the prior art of the invention has not solved how to use these in fields outside the aerospace propulsion.

EXPLANATION OF THE INVENTION

The object of the present invention is a rotor disk as a tangential flow gas turbine and its application as a motor or turbomachine, which overcomes the drawbacks.

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targeted, developing what is described below and is reflected in its essentiality in the first revindication.

The invention is essentially described as a rotor disk, which develops and transmits a work that is obtained from the propulsion generated by the acceleration of the gases of a combustion sustained inside, therefore, based on the foregoing, this device can be qualified, as a turbine, based on the directionality of the propellant flow vector that rotates it, tangential flow and based on its functional set, a non-positive displacement motor, of the internal combustion turbomachine type.

Compared with the current turbomotor, turbojet and turboprop technology, the main advantage of the invention is its greater constructive simplicity, reduced size and weight, where the current technique based on several compressor stages and several turbine stages is more complex to integrate. , of greater size in relation to the power supplied, and of little economic viability in applications outside specific aerospace projects, which cause the industry to opt for other types of engines; mainly plunger motors, where the invention still has a technical advantage, with a lower weight, requires less rotating elements subject to wear, maintenance and lubrication; It has fewer elements subject to high temperatures and bases its operation on a thermodynamically more efficient system, innovating with the use of multiple rotating nozzles that the technique does not contemplate so far.

It should be borne in mind that the purpose of this explanation is to describe the operation of the invention, so that elements such as lubrication system, shaft seal, starting system, controls and electronics applied to turbomachines are obviated. The current technique is broadly contemplated without being of particular importance to the understanding, nor part of the present invention, so that the prior art status will be considered applied in this description. It should be noted that multiple variants contained in this description are derived from the invention, which although they retain the same operating principle, adapt it to different operating environments, as well as different types of fuels / oxidants and different oxidant states. Likewise, as in any other turbine, the invention cannot be understood, or put into practice as an isolated element and depends on other elements that make it work together as a turbomaquine, therefore, on the one hand the isolated invention is described as a turbine of tangential flow gas, and on the other the integration of the same with other components such as turbomachinery and its variants.

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The tangential flow gas turbine, hereinafter, rotor disk (1) is shown in figures 1, 2, 3, 12 and 13, these show the invention to be put into operation with liquid fuels such as kerosene, gasoline, ethanol , methanol and pressurized propane, as well as a gaseous oxidant, such as compressed atmospheric air (21% oxygen) or gaseous oxygen from the evaporation of a source of liquid oxygen by any of the methods that the present technique broadly comprises. It consists of a metal alloy disk or other materials suitable for the particular thermo-mechanical operating requirements; the rotation axis is machined forming the rotation axis hole (2), so that it can be integrated into the axis that hydraulically connects to the radial ducts (5); where these in turn hydraulically connect with the nozzle assemblies (3) having the function of channeling the fuel that comes from the rotation shaft hole to the fuel inlets (6); at the same time, the milling of the oxidant connection to the nozzle assembly (4), which are communicated in this way with the face of the rotor disk exposed to the pressurized oxidant, are practiced in the rotor disk, although these millings admit a multitude of shapes, positions and fluid dynamic designs as seen in figures 3 and 13.

The key component of the tangential flow gas turbine is the nozzle assemblies (3), shown in Figures 4, 5, 6 and 7, its function is to create and maintain an appropriate fuel-oxidant mixture, sustain a combustion and accelerate the combustion product gases to produce a thrust that is transferred to the rotor disk (1). Its integration angle is defined by the axis coinciding with the thrust vector of the nozzle (12) and this in turn, corresponds to the axis of symmetry of the nozzle, being fixed on the rotor disk by interference or any other method of mechanical fixation

According to the angle of integration and force vector that you want to obtain from the nozzle assemblies (3), those in which their angle of integration are parallel to the tangent of the rotor disk (1) at the point of tangential integration are called of integration with mmimal radial and / or axial deviations, also those whose angle of integration is parallel to the tangent of the integration point of the rotor disk with mmimal radial deviations, and an axial angle between ± 90 ° with respect to the tangent. Each nozzle assembly that is integrated in the rotor disk has aligned the fuel inlet (6) a and the radial ducts (5), in turn the oxidant inlet holes (10) communicate with the milling of oxidative connection to nozzle assembly (4) of the rotor disk, although in its variant for liquid oxidant, this comes from radial ducts.

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The nozzle assemblies (3) optionally consist of a fuel injection regulation system by the action of the centrifugal force implicit in the rotation of the rotor disk, which consists of a spring (8) and a valve (7), as can be seen in figures 5 and 6, the valve has holes at a certain height based on the centrifugal force produced by the rotor rotation, the weight of the valve and the force countered by the spring (8), assuming the increase in a rotational speed of the rotor disk, they move the valve through the fuel inlet port (6) by opening and closing the passage of fuel that flows into the fuel injection holes (9) as shown in figure 5, in position of maximum injection flow, before starting to close the flow in case of increasing rotor rotation speed; in the same way, figure 6 shows the position of the valve with the rotor disk stopped.

Considering a constant supply of pressurized fuel available in the fuel inlet (6) and injected this into the combustion chamber through several fuel injection holes (9) and a constant supply of pressurized oxidant entering through the inlet holes of oxidizer (10); The operation of the nozzle assemblies (3) is described as follows: After an initial ignition generated by any method of the present technique, such as an ignition torch for turbine engines, which ignites a small amount of fuel injected into the outlet diffuser of compressor (19), this is conducted through the milling of oxidative connection to nozzle assembly (4) to the nozzle assemblies, where once the ignition phase is finished they sustain a combustion in the combustion chamber (11). As a result of the combustion and thermal expansion of the gases in the nozzle assemblies, these are accelerated and expelled by the nozzle (12), which can describe a multitude of convergent or convergent-divergent forms depending on each application and types of oxidant-fuels, said nozzles (12) based on the acceleration of the gases within them, generate a reaction force opposite to this acceleration or thrust with vector aligned to the axis of the nozzle (12), thrust that is transferred by the nozzle assembly to the disk rotor (1) with the vector that the nozzle has been integrated, where there is at least one tangential component, which drives a rotational force in the rotor disk.

The optionality of the fuel regulator system integrated in the nozzle assemblies (3) is raised by the possibility of the final configuration of the turbomachine as "continuously at maximum power", being able to dispense with the spring (8) and the valve (7), tare the diameter of the fuel injection holes (9) and the gaseous oxidizer inlet holes (10). On the other hand, Figures 5 and 6 show the longitudinal section of a nozzle assembly where its only difference lies in the shape of its nozzle (12) and the position of the regulator of

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fuel, this wants to represent the wide variety of shapes that each impKcita nozzle can and should be designed in the nozzle assembly, which responds to the precise characteristics of each integration; on the other hand, the movement of the exposed fuel regulator system is also shown by difference.

The calculation of the forces applicable in each nozzle, is determined considering these with zero initial velocity independently of the rotor speed, like a rocket motor, since all the velocity vectors at the entrance of the flow in the nozzle are perpendicular to the acceleration vector, where Force = dp / dt of the fluid is fulfilled, where "dp" represents the derivative of the linear momentum and "dt" the derivative of time; also the linear momentum is defined by the expression, p = ymv being the product of the mass, velocity and the Lorentz factor "7", which represents in this system the variation of the mass of the propellant flow in each nozzle corresponding to the velocity relative of the system with respect to the speed of light, defined by the expression, \ or where "c" represents the speed of light and "v" the

yl- v2 / c2

relative speed of the nozzle assemblies (3). Although the Lorentz factor is discarded in other types of thrusters, since these are affected in the same way in the whole propelled assembly, in the case of the invention this does not occur, since a reference system, the housing-frame (14) for example, it may have zero velocity and, however, the nozzles will be in curved motion of thousands of meters per second, considering this and fulfilling that the initial velocity value for the fluids at the entrance of the nozzle is zero, the force developed by the nozzle assemblies will be both greater and greater the equivalent linear velocity thereof in proportion to the Lorentz factor, although this occurs at low speeds in infinitesimal proportions, it is appropriate to determine the thrust force of each nozzle assembly with the formula F = ym Ve + (pe - p0) Ae, where "m" represents the mass flow of the exhaust gas, "See" the effective exhaust velocity, "pe" the static pressure in the outlet plane of each to bera, "p0" system pressure (outlet pressure or ambient pressure) and "Ae" flow area in the nozzle outlet plane.

The tangential flow gas turbine object of the integrated invention as the most essential turbomachine is shown in Figure 10; Figure 8 shows the fixed integration between central axis (16) and rotor disk (1), these two figures show the example of how the fuel enters the tangential flow gas turbine, and how the oxidant pressure is increased . For this, the tangential flow gas turbine is fixed to the central axis by mechanical interference or any other method, said axis hydraulically communicates the fuel inlet (15) with the rotation shaft orifice (2); also transfers the mechanical power generated by the rotor disk to the centrifugal compressor (18) and to the possible elements installed in its final striatum, being

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all of the above supported as shown by a fixed housing-frame part (14), which provides mechanical support to the rotating devices, provides a fixed point of fuel input (15) and channels the oxidant in its path towards the rotor disk , integrating in this case the intake diffuser (17), compressor outlet diffuser (19) and a gas outlet diffuser (20). Considering the turbomachine in operation as detailed in the description of the preferred embodiment of the invention (rotating and with combustion in the nozzle assemblies); The turbomotor shown in Figure 10, using as a kerosene fuel and as an atmospheric air oxidizer, its operation is explained as follows:

■ The oxidant enters through the intake diffuser (17) which is absorbed by the centrifugal compressor (18); Due to the speed of rotation of the compressor, it is compressed upon arrival at the compressor outlet diffuser (19), where it flows through the milling of oxidant connection to nozzle assembly (4) and from there to nozzle assemblies (3).

■ The fuel enters through the fuel inlet (15) at ambient pressure or low pressure, being absorbed by a conduit that communicates it with the groove with perforations in the central axis (16) travels through the axis until it reaches the hole rotation axis (2) of the rotor disk (1), where, due to rotation and the corresponding centrifugal force, its pressure is increased during its flow through the radial ducts (5) until reaching the nozzle assemblies (3).

■ Once the fuel and the oxidant are pressurized and available to the nozzle assemblies, the explanation corresponds to the one previously made on the operation of the same.

■ Once the combustion has been carried out in the nozzle assemblies, a mechanical work is obtained on the rotor disk, which is transferred to the spline of the end of the central axis, said mechanical work available in said spline, can be used for a multitude of purposes, such as the mechanical connection to a generator to produce electricity (as a turbomotor), for the drive of propellers (as a turboprop), or the drive of any other mechanical device, although the gases that are evacuated abroad in this case by a gas outlet (20), can be used to produce a thrust of the turbomachine assembly, (as a turbojet) by applying to figure 10 the compatible elements also shown in figure 11, postcombustion chamber (26) and circumferential nozzle (31 ), instead of a gas outlet (20).

One of the main advantages over the current technique in this invention lies in the nozzle assemblies (3), this design allows the mechanical forces to which they are exposed integrated inside the rotor disk to be low, and therefore enable them, for example,

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to be made in various materials, among them, ceramic materials of high thermal resistance but low toughness, materials that until now the blade turbines cannot assume due to the high mechanical requirements. The nozzle assemblies and therefore the tangential flow gas turbine object of the invention as a turbomaquine, have the capacity to operate under volumes of the exclusively stoichiometric fuel-oxidant mixture, without excess air, therefore at a higher temperature than the turbines of blades of the present technique, which use part of the mechanical work obtained from the turbine to be transferred to the compression phase not only to stoichiometrically supply the oxidant for combustion, but by different methods, to keep the blades at stable temperatures, which It translates into a decrease in turbomachinery performance, greater complexity, as well as greater weight and size due, among other factors, to the oversizing of compressors compared to the invention.

Figures 5 and 6 represent the longitudinal section of the nozzle assembly (3), are shown in its application in the tangential turbine as convergent or convergent-divergent nozzles, said option must be considered in each particular nozzle assembly design, since both are applicable to the tangential flow turbine model, although the convergent nozzle is easier to integrate and operate, the convergent-divergent nozzle is the most efficient model, being able to deliver up to 30% more thrust with it mass flow rate with respect to a convergent nozzle, due to the sum of its fluid accelerations both in its throat, both to the sub-system accelerations of the divergent nozzle section, and therefore, with a useful expansion and acceleration vector, if well, to achieve and maintain the acceleration of a constant fluid flow between the throat and the end of the nozzle, high temperatures and combustion speeds are required, due to the need for and maintain an expansion and acceleration of the gas supersonically, in the best case, ideally adiabatic throughout its journey, conditions not easily obtained using air as an oxidant, since the mere presence of nitrogen as an inert gas at 78% reduces temperatures and combustion speeds remarkably. Although it is relatively easy to obtain a sonic flow in the throat of the nozzle, it is not to maintain the acceleration after it is in the supersonic values until its exit during its entire operating regime, since exposing the divergent nozzle section to a supersonic flow Initially unable to reach the nozzle outlet under its regime the fluid would lose speed and therefore deceleration among other devastating effects energetically speaking, however this approach intends to expose that the invention is not only limited to air as an oxidant, nor limits its use in regimes of speed in nozzle of subsonic or transonic flow, but its use is also considered with other propergoles and supersonico flow rates and

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hypersonic. Figure 7 shows a nozzle assembly adjusted for bipropellants Kquidos, this figure is oriented to show the differences with respect to the nozzle assembly with gaseous oxidant, which lie in the practice of the fuel inlet groove (27), the groove of oxidant inlet (28) and reduction of the diameter of the holes inlet of oxidant (10), its longitudinal section being with the same components to those shown in figures 5 and 6, and its operation corresponds equally to the above Regarding the nozzle assemblies.

The turbojet variant of this invention as a turbomachinery is shown in Figure 11, for now avoiding other aspects of this figure and focusing on the postcombustion chamber elements (26) and the circumferential nozzle (31). These elements have optional character of figure 11, and are replaceable by a gas outlet (20) shown in figure 10 and vice versa. These elements are intended to show the application of the tangential flow gas turbine, such as a turbomachine for jet propulsion by circumferential nozzle. Going into detail in the operation of the turbomachine that integrates the tangential flow gas turbine for the jet propulsion by means of a circumferential nozzle, assuming that the turbomachine is in operation as explained in the previous description and of the preferred realization of the invention, starting from the point at which the gases are expelled by the nozzle assemblies (3) in a subsonic regime, therefore with a convergent nozzle shape, with an oversized throat design such that the energy of combustion is not completely extracted for produce a mechanical work of rotation of the rotor disk (1), but the gases after the nozzle outlet are still disposed of energy in the form of pressure and temperature, pouring into a closed envelope as a postcombustion chamber (26) to be accelerated finally in the circumferential nozzle (31) and from there to the outside, producing in this process an axial vector thrust of the turbomachine assembly as turbojet ctor.

The variant as a heat generating equipment, to be used for example in boilers, is based on the application of the turbomachine assembly of figure no. 10, conceptualizing it, as a gas generator, its operation is as described above for this figure as a turbomachine machine, with the particularity that the gas outlet (20) is connected to the boiler's home in replacement of the burner, such as the gases that leave the turbomotor is still very hot, can be used in said boilers or other elements that require a hot fluid to operate, the main advantage of replacing the boiler burner with the invention, that the burner is an electricity dependent device and that consumes a large amount of it, and the invention, once underway, can function autonomously without any other contribution of energy other than fuel, and even produce electrical energy during

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the use of the boiler, connecting a generator to the spline of the central shaft (16).

The operation in bipropellant systems Kquidos is shown in figure 11, the variant of the transport of the fuel along its central axis (16) according to figure 9, the variant of the nozzle assembly externally by figure 7, and internally by figures 5 and 6, Both the oxidant and the fuel for this design are compounds in a liquid state at room temperature and pressure, oxidizing compounds such as dinitrogen tetroxide, hydrogen peroxide, and fuels such as kerosene, methanol and dimethylhydrazine among others. In this variant, the invention logically lacks the elements related to gas compressors, being as an element of pressure increase of the fuel and oxidant, as detailed above for the rotor disk fuel (1), the centrifugal force at which both fluids are subjected during the rotation of the rotor disk, likewise, the housing-frame (14) for liquid bipropellants mechanically supports the rotating assembly, and provides the fixed points of fuel and oxidant entry. The operation of this variant is explained in a simpler way with oxidizing dinitrogen tetroxide and fuel asymmetric dimethylhydrazine, since the mixture of both is hypergolic and inflames itself when both elements come into contact, although they are non-hypergolic, only The ignition can be added as mentioned in the preferred embodiment but at the exit of the nozzle assembly (3) and an initial turn. The regulation of centrifugal fuel is not applied in the subsequent description of the operation, since this system is suitable for operating in aeronautical, aerospace and underwater propulsion of elements that operate at maximum power during the entire operation, it is decided to regulate the mixture of fuel / oxidant through the oxidizer inlet holes (10) and the fuel injection holes (9) with calibrated diameters that guarantee the correct mixing of both elements and stable operation in known fixed variables at maximum power, which can be describe as follows:

■ With the turbomachine system according to figure 11, completely stopped, the oxidant (dinitrogen tetroxide) enters the system at a previous pressure of 200kPa. through the liquid oxidizer inlet (24), under the same pressure conditions, the fuel (asymmetric dimethylhydrazine) enters the system through the liquid fuel inlet (25), from the respective inputs both flow coaxially separated as shown in figure 9, through the grooves and holes of the central shaft (16), until reaching the rotor disk (1) through the rotation shaft hole (2), where both flow through the radial ducts (5) until reaching the nozzle assemblies (3), where the fuel enters through the fuel inlet groove (27),

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passes to the fuel inlet port (6) and from there to the fuel injection holes (9), where it is injected into the combustion chamber (11); at the same time, the oxidant through the radial ducts (5) flows to the oxidant inlet groove (28), from there to the oxidant inlet holes (10) to be injected into the combustion chamber (11) , where it is mixed with the fuel, which immediately reacts, a combustion takes place and, as a result, an acceleration of the gases that, in the same way described above for the nozzle assemblies, propel and rotate the rotor disk, which together with the spin, by centrifugal force, also increases the injection pressure of oxidant and fuel, which raises the injection amount thereof until progressively reaching the maximum power equilibrium point where the resistance to the passage of both fluids through the injectors is not compensated with the increase of the injection pressure due to the rotation of the rotor, once at this point, although the figure shows the turbojet functionality there are two basic options for the use of the invention on or both at the same time, that of transmitting mechanical work through the stria of the central axis (16) to move for example an electric generator or a propeller, and / or use part of the gases resulting from combustion in the nozzle assemblies ( 3) for the purpose of aeronautical / aerospace propulsion through the use of a postcombustion chamber (26) and a circumferential nozzle (31) or other type of nozzle, or through an angle of integration of the nozzle assemblies with tangential-axial component, whose axial component is the one that develops the thrust. Notwithstanding the foregoing, the aforementioned turbomotor / turbojet system is capable of admitting the regulation of the fuel / oxidant by regulators outside this invention, which the current technique comprises extensively and can be integrated into the liquid oxidant inlet (24) and fuel inlet liquid (25).

Figures 14 and 15 show the solidarity or mechanization association in a single piece in the rotor disk (1) of blades with axial arrangement and radial flow integral to rotor disk (33), in order to use them in the phases described above. in the compression phase, or, in any way that the current technique widely comprises for the turbine blades, being moved by the combustion gases of the nozzle assemblies (3), on the other hand, as seen in Figure 15 , the blades in radial arrangement and axial flow integral to the rotor disk (32), can be used among other uses, to obtain propulsion by forming a fan, with the added advantage of the absence of a torque derived from the fan to the turbomotor assembly, since the force that drives the fan is part of it. The solidarity association of the same, or mechanical integration includes the technique

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In this regard, integration of blades and aerodynamic profiles in rotary devices.

Figure 15 shows the invention, whose variations are caused by the modification of the thrust vector of the nozzle assemblies (3) integrating them with tangential-axial angular component in the rotor disk (1) with the congruent modification of the oxidant connection milling to the nozzle assembly (4) as shown in figures 12 and 13, as well as the integration of radial arrangement vanes and axial flow integral with rotor disc (32). The result of this variation in the operation described above for the turbomachine according to figure 10 is similar and valid with the modifications congruent to the previous description, with the exception that in this variant a tangential vector is obtained that makes the rotor disk move (1) and a thrust vector of the turbomachine assembly, which can be used in aeronautical propulsion, which comes from the axial component of each nozzle assembly and from the dynamic interaction of the fan with the surrounding fluid. Put another way, in this variant there is less power available to perform a mechanical work by means of the center shaft (16) in order to be able to perform a propulsion work directly by means of the nozzle assemblies and the rotor disk, without more elements. Likewise, in accordance with the foregoing, it is evident that the integration of nozzle assemblies with a tangential-axial angle is in the same way applicable in bipropellant systems. Likewise, in figure 15, an axial compressor (35), an axial compressor diffuser (36) and a frame-housing (14) adapted to the above are added to the turbomaquine assembly, which provide greater oxidation compression rates to the invention, and therefore, the possibility of developing greater power per unit of nozzle assembly (3).

From this description, the following main advantageous features of the invention as a whole are disclosed as a turbomachine:

• Constructive simplification.

• Very high power / weight-volume ratios.

• Different applicability, such as turbomotor, turbojet, turboprop and gas generator set in the aerospace and energy industry, as well as part of processes for converting fuels into thermal-mechanical energy applicable to many industries.

• Fuel pressurization is done by the rotor disk, no high pressure fuel pumps are needed.

• Few elements exposed to high temperatures.

• Possibility of integrating liquid bipropellant systems without the need for turbo pumps

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of fuel and oxidant.

BRIEF DESCRIPTION OF THE DRAWINGS

To complement the description that is being carried out and in order to help a better understanding of the characteristics of the invention, it is accompanied as an integral part of said description, a set of drawings in which with an illustrative and non-limiting character, what has been represented next:

• Figure 1.- Shows the isometric view of the tangential flow gas turbine.

• Figure 2.- It shows the top view next to the half of the gas turbine with tangential flow.

• Figure 3.- Shows the view of section A-A of figure 2.

• Figure 4.- Shows the side view of a nozzle assembly.

• Figure 5.- Shows the longitudinal sectional view of a nozzle assembly with a convergent divergent type, with fuel regulator at its maximum flow point (rotor turning at maximum torque).

• Figure 6.- Shows the longitudinal sectional view of a nozzle assembly with a convergent type shape, with fuel regulator in stationary rotor position.

• Figure 7.- Shows the side view of a nozzle assembly for liquid bipropellants, whose differences lie in the practice of grooves and the reduction of the diameter of the oxidizer inlet holes.

• Figure 8.- It shows the side view next to the half, of the tangential flow gas turbine integrated in a central axis.

• Figure 9.- It shows the side view next to the half, of the tangential flow gas turbine, as a variant for liquid bipropellants integrated in a central axis.

• Figure 10.- Shows the isometric view of the tangential flow gas turbine and other integration components such as turbomotor or gas generator block whose housing-frame is open in half.

• Figure 11.- It shows the side view of the tangential flow gas turbine, variant for liquid bipropellants and other integration components such as turbojet and turbomotor as well as the section of the housing-frame for liquid bipropellants.

• Figure 12.- Shows the top view next to the semi-view of the tangential flow gas turbine whose nozzle assemblies have an integration angle with a tangential and axial component at 60 °.

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• Figure 13.- Shows the view of section A-A of figure 12.

• Figure 14.- Shows the side-bottom view showing the blade arrangement

integrated into the tangential flow gas turbine.

• Figure 15.- Shows the side view of the vector tangential flow gas turbine

60 ° tangential-axial and other integration components such as turbojet-

turboprop, with aerodynamic profiles arranged radially to the surface of the tangential flow gas turbine, forming a fan, as well as the section of the frame-frame with axial compressor.

Below is a list of the various elements represented in the figures comprising the invention:

1. Rotor disk.

2. Rotation shaft hole.

3. Nozzle sets.

4. Milling of oxidant connection to nozzle assembly.

5. Radial ducts.

6. Fuel inlet.

7. Valve.

8. Pier.

9. Fuel injection holes.

10. Oxidizer inlet holes.

11. Combustion chamber.

12. Nozzle

14. Housing-frame.

15. Fuel inlet.

16. Central axis.

17. Admission diffuser.

18. Centrifugal compressor.

19. Compressor outlet diffuser.

20. Gas outlet.

24. Liquid oxidizer inlet.

25. Liquid fuel inlet.

26. Postcombustion camera.

27. Fuel inlet groove.

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28. Oxidizer inlet groove.

31. Circumferential nozzle.

32. Radial arrangement and axial flow integral with rotor disk.

33. Axial arrangement and radial flow integral with rotor disk.

35. Axial compressor.

36. Axial compressor output diffuser.

PREFERRED EMBODIMENT OF THE INVENTION

Although there are many applicable environments for the invention, aerospace is the one that best takes advantage of its qualities, so it will be integrated as a turbomotor, in an unmanned aerial vehicle, hereinafter UAV, helicopter type, with the following characteristics of use of the invention

Features and performance of integration of the invention, for 132kW power, estimated weight unit 18Kg, maximum takeoff mass or MTOW of UAV 396kg:

• Rotor disk (1) integrated with: 8 nozzle assemblies (3), arranged in effective radius of 100mm with tangential vector, nominal revolutions of the gas generator-rotor at maximum power or NCPmax, 32900 rpm; maximum power or Pmax, 180 kW; maximum torque, 54.4 N / m at 95% Nc Pmax; peripheral linear speed, 344 m / s.

• Nozzle assemblies (3) integrated with: Convergent-divergent nozzle; unit mass flow rate, 0.056Kg / s; effective gas velocity at nozzle outlet, 1200m / s; throat diameter, 4mm; divergent coefficient, 1.36; net thrust, 68N; outlet pressure, 101kPa.

• Fuel and oxidizer: Jet-A1 and atmospheric air entering the unit through the intake diffuser (17) by means of a fairing diffuser in the UAV, compressed by an axial compressor stage (35), compression ratio, 1.4: 1; centrifugal compressor stage compression ratio, 5: 1; absorbed power, 38kW; fuel / air coefficient, 0.03; specific fuel consumption, 48.3kg / h.

• Starter system: Electric, where the splined end of the central shaft (16) is integrated with a main transmission box provided with a generator-starter unit, which transmits the force necessary for the start of the turbine from the electric energy of a battery, which once the engine is running functions as a generator (10kW absorbed power) of electrical energy to supply

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Electrically, the UAV and turbomotor, on the other hand, ignition torches are located in the compressor outlet diffuser (19).

• Fuel regulation: Centrifugal, the revolutions of the motor-generator set to supply constant helicopters and the variable being the aerodynamic load on the rotor, the centrifugal force regulation system is applicable being preset in flight limits limited in height.

• Other systems: Those inherent in the state of the art.

Taking into account the above, figure 10, to which an axial compressor (35) is added according to figure 15, graphically represents what is described below. The integrated operation of the invention is described as a turbomotor, according to its operating variables, commissioning, operation without load, maximum load operation and shutdown from operation, which are described as follows:

■ Start: A battery makes the generator-starter electrically move in the main transmission box of the UAV, which the tangential flow gas turbine receives through the spline of the central shaft (16) by turning the centrifugal compressor (18). ) and axial compressor (35), activating in turn an ignition device of electric spark and pulverized fuel located in the diffuser compressor outlet (19), when the revolutions reach 10% of Nc, the eight integrated nozzle assembly (3) in the rotor disk (1), and whose masses of the valve (7), under the centrifugal force of the rotation of the rotor cause the spring (8) to contract, it opens and allows the flow of fuel from the fuel inlet (15) located in the frame housing (14) passing through the central axis (16), from there to the rotation axis hole (2), passing through the radial ducts (5) to the nozzle assemblies (3) and their inlet of fuel (6) the valve (7) to the injection holes of fuel (9), where it is injected into the combustion chamber (11).

The oxidant (atmospheric air) enters the inlet diffuser (17) from there to the axial compressor (35), gaining pressure after it, passes through the axial compressor diffuser (36) and enters the centrifugal compressor (18), gaining more pressure after this, it passes through the compressor outlet diffuser (19) where the air flow that is directed to the milling of connection to nozzle assembly (4) is partly inflamed by the fuel and the sparks provided by the ignition torches, entering by said milling, small flames enter the combustion chamber (11) of the nozzle assembly (3) through the oxidant inlet holes (10).

An ignition flame established in the combustion chamber (11) is established in the

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nozzle assemblies (3) a continuous combustion that is accelerating the gases that run through its nozzle (12) and therefore generating an increasing thrust as the number of revolutions increases, and therefore increases with it the fuel pressure, which increases the quantity injected, and the pressure and volume of air supplied by the compressors, thus increasing the revolutions of the rotor disk (1) until the power generated by the tangential flow gas turbine exceeds the power absorbed by the compressor and accessories <60% Nc pmax, where the starter and ignition are deactivated, continuing the starting process up to 110% NC Pmax, where the valve (7) reaches its upper throttle point (figure 5), thus maintaining a speed angular in the constant rotor of ground idle speed.

■ Operation without load: the idle speed at 110% Nc Pmax is considered

■ Operation with load (flight): Starting from idle on the ground, considering the rotors of the UAV at angular speed dividing the angular speed of the tangential rotor of the turbomotor assembly, where the UAV rotors vary their angle of passage and produce a dependently dependent on The mechanical power that the invention generates and transmits through the central axis (16), assuming MTOW, the hovering represents 90% of the maximum power of the turbomotor or PMax and 100% PMax the rise of the helicopter. The invention behaves in such a way that being on the ground in idle speed (obviating the power absorbed by the rotor and accessories on the ground), by varying the rotor pitch of the UAV to produce a stationary flight, it will absorb power from the turbomotor assembly, which will lower the Nc causing the valve to open (7), which will allow a greater flow of fuel to the fuel injection holes (9), said fuel supply will raise the temperature of the combustion gases, which will increase the speed of gas output See, in each nozzle assembly (3) increasing therefore the power transmitted to the central axis (16) until its balance of powers, around 102% Nc Pmax, once in flight, the maximum power is required to the rotor of the invention, which lower the revolutions to 100% Nc Pmax and the valve (7) therefore, slips and allows more fuel flow, delivering the maximum power, assuming a higher power requirement, the maximum torque is 95% Nc Pmax, coinciding with the maximum opening of the valve (7) where the concept of power stability and operation of the invention is confirmed.

■ Shutdown from operation: The fuel supply is cut from (11).

It is further verified that this preferred embodiment is equally achievable in other types of aircraft, electric generators, aircraft apu, land motor units and regenerative boilers among others.

Describing sufficiently the nature of the present invention, as well as the way of putting it into practice, it is stated that, within its essentiality, it may be carried out in other embodiments that differ in detail from that indicated by way of example. , and which will also achieve the protection that is sought, provided that it does not alter, change or modify its fundamental principle.

Claims (27)

5 10 fifteen twenty 25 30 35 1. Non-positive tangential flow displacement motor, characterized in that it comprises: - A housing-frame (14) acting as the frame and stator housing of a compressor assembly for the compression stages of the oxidizing element; - A central axis (16) concentric with the housing, which acts as a rotor of a compressor assembly for the compression and transport stages of the fuel element inside; Y - A cylindrical rotor disk (1) coupled to the central axis in a concentric manner, arranged next to the compression stages, and in which several radial ducts (5) of fuel distribution are arranged to feed corresponding nozzle assemblies (3) located on the periphery of the rotor disk, in which in said nozzle assemblies the fuel and combustion mixture is carried out continuously, and the dynamic acceleration of the gases ejected tangentially to the rotor disk, producing the rotational movement of this and central axis. 2. Non-positive displacement motor of tangential flow according to claim 1, characterized in that each nozzle assembly (3) has: - An oxidizer inlet (10) from the compression stages of the compressor assembly; - A fuel inlet (6) from the corresponding radial duct; - A combustion chamber (11); Y - A nozzle (12) arranged tangentially to the rotor disk. 3. Non-positive tangential flow displacement motor according to claim 2, characterized in that after the fuel inlet (6) a centrifugal fuel inlet valve (7) is integrated, so that the opening and closing of the shutter is regulates by the centrifugal force produced by the rotation of the rotor disk. 4. Non-positive displacement motor of tangential flow according to claim 3, characterized in that the intake valve has a control spring (8) that compresses against the centrifugal force. 5. Non-positive displacement motor of tangential flow according to claim 2, characterized in that the nozzle assembly has a cylindrical symmetry with the axis of symmetry arranged tangentially to the turbine disc assembly. 5 10 fifteen twenty 25 30 35 6. Non-positive displacement motor of tangential flow according to claim 2, characterized in that the outlet nozzle is of convergent type. 7. Non-positive displacement motor of tangential flow according to claim 2, characterized in that the outlet nozzle is of convergent-divergent type. 8. Non-positive tangential flow displacement motor according to claim 1, characterized in that the compression stages of the compressor assembly include at least one axial compressor. 9. Non-positive tangential flow displacement motor according to claim 1, characterized in that the compression stages of the compressor assembly include at least one centrifugal compressor. 10. Non-positive displacement motor of tangential flow according to claim 1, characterized in that the central axis has a geared end to transmit the mechanical turning force.
11. Non-positive tangential flow displacement motor according to claim 2, characterized in that some of the nozzles are located according to a mixed tangential-axial vector, providing, in addition to the rotary movement of the rotor disc assembly and the central axis, an axial propulsion according to the central axis.
12. Non-positive displacement motor of tangential flow according to claim 1, characterized in that the central axis is coupled to an electric starter motor.
13. Tangential flow non-positive displacement motor according to claim 1, characterized in that the central axis is coupled to an electric generator.
14. Non-positive tangential flow motor according to claim 1, characterized in that the housing-frame directs the burnt gases from the nozzle assemblies in an axial direction according to the central axis by means of post-bushing chambers and a concentric circumferential nozzle with said axis.
15. Non-positive tangential flow displacement motor according to claim 1, 5 10 fifteen twenty 25 30 35 characterized in that at the outlet of the rotor disk there is a blade turbine (32, 33) integral to the rotor disk, which is moved by the burnt gases from the nozzle assemblies.
16. Non-positive tangential flow displacement motor according to claim 15, characterized in that said turbine is formed by axial arrangement vanes and radial flow (33).
17. Tangential flow non-positive displacement motor according to claim 15, characterized in that the turbine is formed by blades of radial arrangement and axial flow (32).
18. Non-positive tangential flow displacement motor according to claim 1, characterized in that said non-positive displacement motor is coupled to a mechanism for its operation as a turbomotor.
19. Tangential flow non-positive displacement motor according to claim 1, characterized in that said non-positive displacement motor is coupled to a propeller or propeller for operation as a turboprop or turbofan. 20. Tangential flow non-positive displacement motor according to claims 11 or 14, characterized in that said non-positive displacement motor operates as a turbojet. 21. Non-positive tangential flow displacement motor according to claim 1, characterized by: - The central shaft (16) transports the fuel and oxidant (oxidizer) elements separately, without prior compression steps of the oxidant. - The rotor disc has pairs of parallel radial ducts (5) for the separate distribution of the fuel and oxidant elements to each nozzle assembly. 22. Tangential flow non-positive displacement motor according to claim 19, characterized in that the turbofan is coupled directly to the rotor disk by means of radially arranged vanes and axial flow integral to the rotor disk (32). 23. Non-positive tangential displacement motor according to claim 13, characterized in that the electric generator coupled to the shaft also acts as a motor for 5 10 fifteen twenty 25 30 electric start during the boot process. 24. Tangential flow non-positive displacement motor according to claim 1, characterized in that said non-positive displacement motor functions as heat generating equipment for boilers. 25. Method of operation of the non-positive tangential flow displacement motor defined in claim 1, characterized in that it comprises the steps of: - Continuous intake of oxidizer at the entrance of the compression stages; - Continuous compression of the oxidizer in the compression stages; - Continuous combustion of the fuel-combustion mixture in the combustion chamber (11) of each nozzle assembly; Y - Ejection of the gases burned by the nozzle (12) of the outlet of each nozzle assembly. 26. Method of operation of the non-positive displacement motor according to claim 17, characterized in that the nozzle assemblies (3) have the ability to operate under stoichiometric ratios of the fuel-oxidizing mixture. 27. Starting method of the non-positive tangential flow displacement motor defined in claims 12 and 13, characterized in that it comprises the steps of: - Start-up of the starter motor and coupling to the central shaft (16), so that the compressor assembly and the rotor disk are rotated. - Opening by the centrifugal force of the fuel inlet valves (7), which will allow the fuel to be mixed with the oxidizer from the compression stages of the compressor assembly. - Activation of the ignition device, which ignites a combustible mixture at the exit of the last compression stage. - Transmission downstream of the combustion to the combustion chambers (11) of the nozzle assemblies, thereby accelerating the exit of the burnt gases and the rotation of the central shaft and the compressor. - Deactivation of the ignition device and the starter motor when the power generated by the non-positive displacement motor exceeds the power absorbed by the compressor assembly.