ES2691990B1 - Non-positive tangential flow displacement motor - Google Patents

Description

D E S C R I P C I Ó N

MOTOR OF NON POSITIVE DISPLACEMENT OF TANGENTIAL FLOW

SECTOR OF THE TECHNIQUE

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, electric power generation, propulsion of land vehicles and internal combustion machinery.

BACKGROUND OF THE INVENTION

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 ^ initial temperature of the fluid and T2 the temperature

end 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 from 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 excess energy is used according to the mentioned types, the turbomotors use more turbine stages to transmit said mechanical energy to a device, the turboprops similarly to a propeller type propeller or fan, and turboreactors to accelerate the fluid in a nozzle with the remaining kinetic energy in order to produce a thrust.

In this definition, all processes in which their energy balance is negative, that is, require energy for their development, are named as an energy deficit, a fact that is vital 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 blade temperatures to acceptable values, both by lowering the temperature of the primary combustion flow, as flow secondary to cool the blades directly, which ultimately results in a series of p both mechanical and thermodynamic losses, or what is the same, a notable reduction in efficiency, in addition to a construction set of greater volume and 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 supersonic-hypersonic regime are some of the most efficient thermodynamic machines and that are closer to the maximum Carnot performance, the prior art 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. pointed, developing what is described below and is reflected in its essentiality in the first claim.

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 sustained combustion 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 engine, of the internal combustion turbomachinery type.

Compared to the current technique of turbomotors, turbo-reactors and turboprops, 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 smaller, precise weight of 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 prior art will be deemed applied in this description. It should be noted that multiple variants contained in this description emerge 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 turbomachine, therefore, on the one hand, the isolated invention is described as a turbine. of tangential flow gas, and on the other the integration of it with other components such as turbomachinery and its variants.

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 , pressurized methanol and 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 an alloy metal disk or other materials suitable for the particular thermo-mechanical operating requirements; the rotation axis is machined forming the rotation axis hole (2), such that an axis that connects hydraulically to the radial ducts (5) can be integrated therein; where these in turn connect hydraulically with the nozzle assemblies (3) with the function of channeling the fuel that comes from the axis of rotation orifice 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 oxidant under pressure, 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, their 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 one wishes to obtain from the nozzle assemblies (3), those in which their integration angle is parallel to the tangent of the rotor disk (1) at the point of tangential integration are called of integration with minimum 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 minimum 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, it comes from radial ducts.

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 is composed 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 that based on the centrifugal force produced by the rotor rotation, the weight of the valve and the force counteracted by the spring (8), assuming the increase in a rotational speed of the rotor disk, 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 increased rotation speed of the rotor; 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 by means of several fuel injection holes (9) and a constant supply of pressurized oxidant entering through the inlet holes 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 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 turbomachinery as "continuously at maximum power", and the spring (8) and the valve (7) can be dispensed with, tare the diameter of the fuel injection holes (9) and the gas oxidant 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 fuel, this wants to represent the wide variety of forms that each implicit 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 regulatory system of the exposed fuel 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 "y", 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 y, l - \ v 2 / c o2 where "c" represents the speed of light and "v" the relative speed of the nozzle assemblies (3) . Although the Lorentz factor is discarded in other types of propellants, 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 can have zero velocity and yet the nozzles will be in curvilinear 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, " Ve" the effective exhaust velocity, "pe" the static pressure in the outlet plane of each nozzle, "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 invention integrated 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); It 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 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 entry of fuel (15) and channels that of the oxidant on its way to 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 effect of 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 oxidative connection to nozzle assembly (4) and from there to the 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 their operation.

■ Once the combustion has been carried out in the nozzle assemblies, a mechanical work is obtained in 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 (such as a turbomotor), for the drive of propellers (such 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, afterburner (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, to be made in various materials, including ceramic materials of high thermal resistance but low toughness, materials that until now the turbines of blades can not assume due to the high mechanical requirements. The nozzle assemblies and therefore the tangential flow gas turbine object of the invention as a turbomachine, have the ability 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 the compressors in comparison 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 the supersonic 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 supersonic flow between the throat and end of the nozzle, high temperatures and combustion speeds are required, due It is the need to maintain a gas expansion and acceleration supersonic, at best, ideally adiabatic throughout its journey, conditions not easily obtained using air as an oxidizer, since the mere presence of nitrogen as 78% inert gas reduces combustion temperatures and speeds significantly. 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 in 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 supersonic regime the fluid would lose speed and therefore deceleration among other devastating effects energetically speaking, however this approach aims to expose that the invention is not only limited to air as an oxidizer, nor limits its use in regimens of speed in subsonic or transonic flow nozzle, but its use is also considered with other propergoles and supersonic flow rates and ersonic hip. An adjusted nozzle assembly for liquid bipropellants is shown in figure 7, 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 the 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 post-combustion chamber elements (26) and the circumferential nozzle (31). These elements are optional in Figure 11, and are replaceable with 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, as a turbomachine for the jet propulsion by circumferential nozzle. Going into detail in the operation of the turbomachinery that integrates the tangential flow gas turbine for 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 embodiment 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 combustion energy 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 post-combustion 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 turbo assembly machine as turbojet.

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 instead of the burner, 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 by the invention, that the burner is an electricity dependent device and that it consumes a large amount of it, and the invention, once underway, can operate autonomously without any other energy input other than fuel, and even produce electrical energy during the use of the boiler, connecting a generator to the spline of the central shaft (16).

The operation in liquid bipropellant systems 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 increasing the pressure 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 bi-propellants 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 the 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 whole operation, the regulation of the mixture of fuel / oxidant by means of the oxidizer inlet holes (10) and the fuel injection holes (9) being made 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), it 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, 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 amount of injection 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 in injection pressure due to rotor rotation, once at this point, although the figure shows the turbojet functionality there are two basic options for use of the invention or both at the same time, of transmitting mechanical work through the spline 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 assemblies nozzle (3) for aeronautical / aerospace propulsion purposes by using a post-combustion chamber (26) and a circumferential nozzle (31) or other type of nozzle, or through an angle of integration of the nozzle assemblies with tangential component- axial, whose axial component is what 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 association of solidarity or mechanization 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 broadly 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, for obtaining 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 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) by 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 blades and axial flow integral with rotor disk (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 central shaft (16) spline in order to be able to perform a propulsion work directly by means of the nozzle assemblies and the rotor disk, without more elements. Also, in accordance with the foregoing, it is evident that the integration of nozzle assemblies with an axial-tangential angle is similarly 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 turbomachinery, which provide greater oxidation compression ratios 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 of fuel and oxidant.

BRIEF DESCRIPTION OF THE DRAWINGS

To complement the description that is being made and in order to help a better understanding of the characteristics of the invention, a set of drawings is attached as an integral part of said description, where illustrative and non-limiting nature 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 section view of a nozzle assembly with a convergent divergent type, with fuel regulator at its maximum flow point (rotor rotating at maximum torque).

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

• Figure 7.- It 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 bi-propellants 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 is open in half.

• Figure 11.- 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.- It shows the top view next to the half view of the tangential flow gas turbine whose nozzle assemblies have an integration angle with a tangential and axial component at 60 °.

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

• Figure 14.- Shows the side-bottom view showing the arrangement of integral blades integrated in the tangential flow gas turbine.

• Figure 15.- It shows the side view of the tangential-flow gas turbine of 60 ° tangential-axial vector 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 housing housing 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. Post combustion chamber.

27. Fuel inlet groove.

28. Oxidizer inlet groove.

31. Circumferential nozzle.

32. Radial arrangement blade and axial flow integral with rotor disc.

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

35. Axial compressor.

36. Axial compressor output diffuser.

PREFERRED EMBODIMENT OF THE INVENTION

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

Characteristics and integration performance 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 100mm effective radius 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.

• Starting 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 power of a battery, which once the engine is running, functions as a generator (10kW absorbed power) of electrical energy to supply 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 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) according to figure 15 is added, graphically represents what is described below. The integrated operation of the invention as a turbomotor is explained, 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 N c, the eight nozzle assembly (3) integrated 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 inlet of fuel (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 fuel inlet (6) the valve (7) towards the holes of fuel injection (9), where it is injected into the combustion chamber (11).

The oxidant (atmospheric air) enters through 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 oxidizer inlet holes (10).

An ignition flame established in the combustion chamber (11) is established in the nozzle assemblies (3) a continuous combustion that accelerates 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 rotors of the UAV vary their angle of passage and produce a lift depending 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 ascent of the helicopter. The invention behaves in such a way that while on the ground in idle mode (bypassing 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 opening of the valve (7), which will allow a greater flow of fuel to the fuel injection orifices (9), said fuel supply will raise the temperature of the flue gases, which will increase the speed gas outlet See, in each nozzle assembly (3) increasing therefore the power transmitted to the central axis (16) until its power balance, around 102% Nc Pmax, once in flight, the rotor is required the maximum power of the invention, which lower the revolutions to 100% Nc Pmax and the valve (7) therefore, slides and allows more fuel flow, the maximum power being delivered, assuming a power requirement at higher, 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 realizable in other types of aircraft, electric generators, aircraft apu's, 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 noted that, within its essentiality, it may be implemented in other embodiments that differ in detail from that indicated by way of example. , and which will also achieve the protection sought, provided that it does not alter, change or modify its fundamental principle.

Claims (27)

1. Non-positive tangential flow displacement motor, characterized in that it comprises: - A frame housing (14) acting as a 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 disc (1) coupled to the central axis concentrically, arranged after the compression stages, and in which several radial fuel distribution ducts (5) are arranged to feed corresponding nozzle assemblies (3) located on the periphery of the rotor disk, in which in said nozzle assemblies the mixing of fuel and combustion is carried out continuously, and the dynamic acceleration of the gases ejected tangentially to the rotor disk, producing the rotational movement of the latter and central axis. 2. Non-positive tangential flow displacement motor 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 type 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 tangential flow displacement motor according to claim 3, characterized in that the intake valve has a control spring (8) that works under compression against the centrifugal force. 5. Non-positive tangential flow displacement motor 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. 6. Tangential flow non-positive displacement motor according to claim 2, characterized in that the outlet nozzle is of convergent type. 7. Non-positive tangential flow displacement motor 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 tangential flow displacement motor according to claim 1, characterized in that the central axis has a geared end for transmitting the mechanical turning force. 11. Non-positive tangential flow displacement motor according to claim 2, characterized in that some of the nozzles are positioned 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 tangential flow displacement motor according to claim 1, characterized in that the central axis is coupled to an electric starter motor. 13. Non-positive tangential flow displacement motor according to claim 1, characterized in that the central axis is coupled to an electric generator. 14. Non-positive tangential flow displacement motor according to claim 1, characterized in that the frame housing directs the burned gases from the nozzle assemblies axially along the central axis by means of post combustion chambers and a concentric nozzle concentric with the said axis. 15. Tangential flow non-positive displacement motor according to claim 1, 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 displacement motor according to claim 15, characterized in that said turbine is formed by axially arranged blades and radial flow (33). 17. Tangential flow non-positive displacement motor according to claim 15, characterized in that the turbine is formed by radial arrangement blades and axial flow (32). 18. Tangential flow non-positive 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 in that: - The central shaft (16) transports the fuel and oxidant (oxidizing) elements separately, without prior compression stages 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. Non-positive tangential flow displacement motor according to claim 19, characterized in that the turbofan is directly coupled to the rotor disk by means of radially arranged blades and axial flow integral to a rotor disk (32). 23. Non-positive tangential flow displacement motor according to claim 13, characterized in that the electric generator coupled to the shaft also acts as a motor of 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. Non-positive displacement motor operating method 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.