Supersonic Inlet With Zero Cowl Angle And Bypass Splitter
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Solution Overview
Problem
Conventional supersonic inlet designs face challenges in reducing sonic boom signature and drag, as they often result in complex geometries that increase drag and reduce pressure recovery at high supersonic speeds, and traditional methods to control drag either lead to integration issues or compromise pressure recovery.
Innovation Solution
A supersonic inlet design featuring a relaxed isentropic compression surface with a zero or near-zero cowl lip angle and a bypass splitter to capture initial and terminal shocks internally, reducing cowl lip angle and drag while maintaining pressure recovery, and employing a nacelle bypass to isolate and dispose of flow distortions.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Loss of energy
If traditional shock-on-lip focusing is used to maximize pressure recovery, then pressure recovery is improved, but inlet drag increases due to larger cowl lip angle required at high supersonic speeds
Solution Approach 1:
The inlet flow is segmented into two separate paths: a primary flow path that passes through the compression surface and terminal shock, and a bypass flow path that goes around the compression surface. This segmentation allows the bypass flow to avoid the strong shocks and flow turning, reducing drag while the primary path maintains pressure recovery.
Solution Approach 2:
A bypass splitter acts as an intermediary component that divides the incoming supersonic flow into primary and bypass streams. The bypass splitter enables the bypass flow to follow a different trajectory that avoids the adverse effects of shock-on-lip focusing, thereby reducing drag without compromising the primary flow's pressure recovery.
2Ease of operation
If cowl lip angle is increased to align with local flow at higher supersonic speeds, then flow alignment is improved, but inlet frontal area and drag increase
Solution Approach 1:
The flow alignment problem is resolved by segmenting the flow into bypass and primary paths. The bypass flow follows a different trajectory with smaller flow turning requirements, allowing the cowl lip angle to be reduced while still achieving proper flow alignment through the compression surface for the primary stream.
3Force
If complex 3-D inlet geometry is used to control drag, then drag is reduced, but aircraft integration becomes difficult and flow distortion increases
Solution Approach 1:
The inlet is segmented into simple geometric components (centerbody, cowl, bypass splitter) that can be easily integrated into the aircraft structure. The bypass flow path provides drag control without requiring complex 3-D geometry, thus avoiding integration difficulties while still achieving the desired drag reduction.
4Force
If terminal shock Mach number is increased to reduce flow turn angle, then drag is reduced, but pressure recovery decreases due to stronger terminal shock
Solution Approach 1:
The segmentation of flow into bypass and primary paths allows the primary flow to maintain a moderate terminal shock Mach number for good pressure recovery, while the bypass flow contributes to drag reduction. The overall system achieves both goals simultaneously through flow segmentation rather than relying on extreme terminal shock conditions.
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 significantly reduces the sonic boom signature and drag by minimizing cowl shock strength and flow spillage, while maintaining efficient pressure recovery and engine performance by isolating flow distortions and routing excess airflow effectively.
Implementation Method 1
The compression surface may be configured to generate a second shock wave
Implementation Method 2
A measurement of inlet operation efficiency is the total pressure lost in the air stream between the entrance side and the discharge side of the inlet
Implementation Method 3
A diffuser further decelerates the resulting flow from the strong terminal shock to a speed corresponding to the requirements of the gas turbine engine
Data Source
AI summary
A supersonic inlet having a cowl lip configured to capture the conic shock and exhibit a zero or substantially zero cowl angle is disclosed herein. The inlet may be configured to employ a relaxed isentropic compression surface and an internal bypass. The nacelle bypass may prevent flow distortions, introduced by the capture of the conic shock, from reaching the turbomachinery, thereby allowing the cowl angle to be reduced to zero or substantially zero. Such a cowl angle may reduce the inlet's contribution to the overall sonic boom signature for a supersonic aircraft while allowing for an increase in engine pressure recovery and a subsequent improvement in generated thrust by the engine.