Phased composite cooling structure and method for aircraft rudder

Abstract

The present invention discloses a staged composite cooling structure for an aircraft rudder, comprising a pressure pump, a rudder disposed at the airflow output end of the pressure pump, a microchannel disposed within the rudder near the interface of the pressure pump, an internal space disposed at the end of the microchannel and within the rudder, a rib disposed within the inner cavity of the internal space and near the end of the microchannel, and a plurality of air film holes disposed on the supporting end surface of the rudder, wherein the plurality of air film holes connect the interior and exterior of the rudder. Advantages: This invention is highly innovative in applying composite cooling technology to rudder design. Furthermore, when a rudder incorporating this composite cooling technology is introduced, the maximum temperature of the corresponding structure can be reduced when the aircraft's flight Mach number is greater than 6, thereby improving the reliability and service life of the aircraft's rudder.

Description

Technical Field

[0001] The present invention relates to a staged composite cooling structure and method for an aircraft rudder. When the aircraft's flight Mach number is greater than 6, the equipment can implement a composite cooling process for aircraft components that are subjected to extremely high heat flux density through aerodynamic heating, thereby lowering the maximum temperature of the corresponding components and reducing the temperature gradient, thereby reducing the thermal stress of the components, improving the reliability and service life of the corresponding components, and ensuring the aerodynamic performance and maneuverability of the aircraft. Background Art

[0002] When an aircraft flies at a Mach number greater than 6, aerodynamic heating subjects protruding components like the rudder to extremely high heat fluxes, with surface temperatures reaching 2000K-3900K. Existing materials are unable to maintain good performance and stability at such high temperatures.

[0003] The cooling technologies commonly used at present include the following:

[0004] Divergent cooling: Divergent cooling, when the same technology is applied to the same model, may result in overcooling of low-temperature areas and insufficient cooling of high-temperature areas.

[0005] Air cooling technology: Air cooling technology, there is no cooling air source during high-speed flight and the air contains dust, which can easily cause dust to block the channel. Summary of the Invention

[0006] The purpose of the present invention is to provide a staged composite cooling structure and method for an aircraft rudder. In the case that the existing materials used for rudders are difficult to withstand ultra-high temperatures for a long time, a high-efficiency staged composite cooling technology is used (first, an appropriate liquid cooling medium is selected to utilize the heat absorption capacity of the cooling medium to absorb heat through phase change boiling heat exchange in a microchannel, then a gaseous cooling medium is utilized to flow in the complex channels inside the component to absorb heat, and finally the gaseous cooling medium flows out along the air film holes to form a uniform air film to achieve external cooling of the component) to provide thermal protection for the rudder of a hypersonic aircraft, so as to solve the problems raised in the above-mentioned background technology.

[0007] To achieve the above-mentioned objectives, the present invention provides the following technical solutions: a staged composite cooling structure for an aircraft rudder, comprising a pressure pump, a rudder being provided at the airflow output end of the pressure pump, a microchannel being provided in the rudder near the interface of the pressure pump, an internal space being provided at the end of the microchannel and inside the rudder, a rib being provided in the inner cavity of the internal space and near the end of the microchannel, a plurality of air film holes being provided on the supporting end surface of the rudder, and the plurality of air film holes being connected to the inside and outside of the rudder.

[0008] A cooling method using a staged composite cooling structure for a hypersonic vehicle rudder, the method comprising the following steps:

[0009] Step 1: pumping in liquid cooling medium;

[0010] Step 2: Cooling medium flows into the microchannel for cooling;

[0011] Step 3: The cooling medium enters the rudder for cooling;

[0012] Step 4: The gaseous cooling medium reaches the surface of the rudder through the air film holes to form a protective air film.

[0013] Preferably, the liquid cooling medium is pumped into the storage space via a pressure pump and flows into the rudder via a rudder shaft interface.

[0014] Preferably, the cooling medium flows into the microchannel for cooling, and the microchannel is set up so that the cooling medium can undergo phase change boiling heat exchange inside the rudder to absorb heat and finally vaporize into gaseous cooling medium.

[0015] Preferably, the cooling medium enters the interior of the rudder for cooling in the form of a gas after leaving the microchannel, and can flow through the internal cooling channel for heat exchange, using the ribs inside to enhance heat transfer, thereby completing further cooling of the rudder.

[0016] Preferably, the gaseous cooling medium reaches the outer surface of the rudder through the air film holes to form a protective air film. This is achieved by opening air film holes on the surface of the rudder so that the gaseous cooling medium can flow out and form a protective air film on the outer surface of the rudder to cool the outer surface of the rudder.

[0017] Compared with the prior art, the present invention has the following beneficial effects:

[0018] The composite cooling technology method of the present invention consists of three cooling steps: utilizing the heat absorption capacity of the cooling medium to absorb heat through phase change boiling heat exchange in the microchannel, the gaseous cooling medium flowing in the complex channels inside the component to absorb heat and enhance heat transfer, and the gaseous cooling medium flowing out along the air film holes to form a uniform air film to achieve external cooling of the component. The invention is highly innovative in applying the composite cooling technology method to rudder design, and after introducing a rudder with this composite cooling technology, when the flight Mach number of the aircraft is greater than 6, the maximum temperature of the corresponding structure can be reduced, thereby improving the reliability and service life of the aircraft rudder. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 A schematic diagram of a staged composite cooling structure for an aircraft rudder provided by the present invention;

[0020] Figure 2 This is a flow chart of the method adopted by the staged composite cooling technology for aircraft rudders provided by the present invention.

[0021] In the figure: 1. Pressure pump; 2. Rudder; 3. Microchannel; 4. Internal space; 5. Rib; 6. Air film hole. DETAILED DESCRIPTION

[0022] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0023] See also Figure 1-2 The present invention provides a technical solution: a staged composite cooling structure for an aircraft rudder, comprising a pressure pump 1, the airflow output end of the pressure pump 1 being connected to a rudder 2 through a rudder shaft interface, a microchannel 3 being installed in the rudder 2 near the interface of the pressure pump 1, an internal space 4 being installed at the end of the microchannel 3 and inside the rudder 2, a rib 5 being installed in the inner cavity of the internal space 4 and near the end of the microchannel 3, a plurality of air film holes 6 being provided through an opening on the support end surface of the rudder 2, and the plurality of air film holes 6 being connected to the inside and outside of the rudder 2.

[0024] A cooling method using a staged composite cooling structure for a hypersonic vehicle rudder, the method comprising the following steps:

[0025] Step 1: pumping in liquid cooling medium;

[0026] Step 2: Cooling medium flows into microchannel 3 for cooling;

[0027] Step 3: The cooling medium enters the interior of the rudder 2 for cooling;

[0028] Step 4: The gaseous cooling medium reaches the surface of the rudder 2 through the air film holes 6 to form a protective air film.

[0029] The liquid cooling medium is pumped into the storage space via the pressure pump 1 and flows into the rudder 2 via the rudder shaft interface.

[0030] The cooling medium flows into the microchannel 3 for cooling. The microchannel 3 is set up so that the cooling medium can undergo phase change, boiling heat exchange, and absorb heat inside the rudder 2, and finally vaporize into a gaseous cooling medium.

[0031] The cooling medium enters the rudder 2 for cooling. After leaving the microchannel 3, the cooling medium enters the internal space 4 in gaseous form and can flow through the internal cooling channel for heat exchange. The ribs 5 inside are used to enhance heat transfer and further cool the rudder 2.

[0032] The gaseous cooling medium reaches the surface of the rudder 2 through the air film holes 6 to form a protective air film. The air film holes 6 are opened on the surface of the rudder 2 so that the gaseous cooling medium can flow out and form a protective air film on the outer surface of the rudder 2 to cool the outer surface of the rudder 2.

[0033] Working principle: When the aircraft is used for hypersonic flight and the surface temperature of the rudder 2 component is too high due to the influence of aerodynamic heating, the pressure pump 1 can be started to pump the liquid cooling medium stored inside the aircraft into the rudder 2 through the rudder shaft interface. The liquid cooling medium pumped into the rudder 2 is first transported to its internal microchannel 3, which allows the liquid cooling medium to quickly absorb heat through the three processes of "unsaturated liquid-gas-liquid two-phase flow-superheated gas" and vaporize into gaseous cooling medium for subsequent internal and external cooling. Then, after the liquid cooling medium stored inside the aircraft is converted into a gaseous cooling medium, efficient convection can be carried out through the complex channels inside the rudder 2 component to further enable the cooling medium to absorb heat and transfer heat, thereby achieving cooling, and the rib 5 structure is used in the inner cavity of the internal space 4 to enhance heat transfer. Finally, when the gaseous cooling medium reaches the air film hole 6 of the rudder 2, it will flow out along the air film hole 6 to form a uniform air film. At this time, the protective air film formed by the gaseous cooling medium on the outside of the rudder 2 achieves further cooling of the rudder 2 to achieve thermal protection characteristics.

[0034] It should be noted that, in this document, relational terms such as first and second, etc., are used only to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprises," "comprising," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that includes a list of elements includes not only those elements but also other elements not explicitly listed, or elements inherent to such process, method, article, or apparatus.

[0035] While embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions, and variations may be made to these embodiments without departing from the principles and spirit of the invention, and that the scope of the invention is defined by the appended claims and their equivalents.

Claims

1. A phased composite cooling method for an aircraft rudder, characterized in that: The method comprises a structure, wherein the structure comprises a pressure pump (1), a rudder (2) is provided at an airflow output end of the pressure pump (1), a microchannel (3) is provided in the rudder (2) near an interface of the pressure pump (1), an internal space (4) is provided at the end of the microchannel (3) and inside the rudder (2), a rib (5) is provided in the inner cavity of the internal space (4) and near the end of the microchannel (3), a supporting end surface of the rudder (2) is provided with a plurality of air film holes (6), and the plurality of air film holes (6) are connected to the inside and outside of the rudder (2); The method is characterized in that: Step 1, pumping in a liquid cooling medium, wherein the liquid cooling medium is pumped in from a storage space via a pressure pump (1) and flows into the rudder (2) via a rudder shaft interface; Step 2: The cooling medium flows into the microchannel (3) for cooling. The cooling medium flows into the microchannel (3) for cooling. The microchannel (3) is set up to allow the cooling medium to undergo phase change boiling heat exchange inside the rudder (2) to absorb heat and eventually vaporize into a gaseous cooling medium. Step 3, the cooling medium enters the interior of the rudder (2) for cooling. The cooling medium enters the interior of the rudder (2) for cooling by allowing the cooling medium to enter the interior space (4) in a gaseous form after leaving the microchannel (3) and to flow and exchange heat through the internal cooling channel. The ribs (5) therein are used to enhance heat transfer, thereby further cooling the rudder (2). In step 4, the gaseous cooling medium reaches the outer surface of the rudder (2) through the air film holes (6) to form a protective air film. The gaseous cooling medium reaches the outer surface of the rudder (2) through the air film holes (6) to form a protective air film. The gaseous cooling medium reaches the outer surface of the rudder (2) through the air film holes (6) to form a protective air film. The air film holes (6) are provided on the surface of the rudder (2) so that the gaseous cooling medium can flow out and form a protective air film on the outer surface of the rudder (2) to achieve cooling of the outer surface of the rudder (2).

Images