Fan Compressor Blade Laminar Flow Control
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Solution Overview
Problem
Conventional fan rotor blades experience high friction loss due to early transition of the boundary layer from laminar to turbulent state, resulting from rapid deceleration of air flow and sharp increase in flow channel area, which reduces the laminar flow region and increases frictional drag.
Innovation Solution
The blade is designed with a subsonic and transonic region, where the blade surface angle change rate is controlled to delay the transition of the boundary layer from laminar to turbulent state by modifying the airfoil shape, specifically by reducing the peak blade surface angle change rate in the subsonic region and increasing it in the transonic region, thereby enlarging the laminar flow region.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If the blade surface angle decreases sharply from leading edge to trailing edge, then the flow channel area increases rapidly, but the boundary layer transitions early from laminar to turbulent state, increasing friction loss
Solution Approach 1:
The patent applies parameter changes by precisely controlling the blade surface angle change rate distribution along the chord direction. By setting specific constraints on the blade surface angle change rate (δ) at different chord positions (x/c), the invention modifies the flow acceleration characteristics to delay boundary layer transition, thereby reducing friction loss while maintaining productivity.
Solution Approach 2:
The patent implements local quality by applying different blade surface angle change rate constraints to different regions of the blade surface. The subsonic region (root side) and transonic region (tip side) have distinct parameter ranges, allowing optimized control of boundary layer transition in each region to minimize overall friction loss.
2Loss of energy
If the blade surface angle change rate is increased to enlarge laminar flow region, then friction loss is reduced, but the flow deceleration control becomes more complex
Solution Approach 1:
The patent resolves the complexity issue by establishing clear quantitative parameter ranges for the blade surface angle change rate at specific chord positions. These parameter constraints provide a systematic design methodology that simplifies the complex blade shaping process while achieving the goal of enlarging laminar flow region and reducing friction loss.
3Loss of energy
If the boundary layer transition is delayed to enlarge laminar flow region, then frictional drag is reduced, but the blade surface angle distribution becomes more difficult to design
Solution Approach 1:
The patent addresses manufacturing precision requirements by defining specific parameter ranges for the blade surface angle change rate at standardized chord positions. These quantified parameters serve as clear manufacturing guidelines, making it easier to achieve the desired blade geometry with appropriate precision while delaying boundary layer transition.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This design effectively delays the boundary layer transition, enlarging the laminar flow region and reducing friction loss by controlling the deceleration of air flow around the blade, resulting in improved efficiency.
Implementation Method 1
the transition of the boundary layer from a laminar state to a turbulent state
Implementation Method 2
enlarging a laminar flow region over the blade surface
Implementation Method 3
reducing friction loss by controlling the deceleration of air flow around the blade
Data Source
AI summary
A blade of a fan or compressor that reduces loss by enlarging a laminar flow region over a blade surface is provided. The blade is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. A blade surface angle change rate is based on an angle formed by a tangent to the blade surface and the axis of the engine, the leading edge blade surface angle, and the trailing edge blade surface angle at. In each of the subsonic region and the transonic region, values of the blade surface angle change rate on the pressure and suction surfaces are defined at predetermined axial locations along the chord on the pressure and suction surfaces.