Combined cycle propulsion shows great potential for wide speed and altitude range flights. However, the jets of different types of engines may interact and form complex flow structures. Given that the specific effects of jets interaction are closely related to combined nozzle configuration and working conditions, this paper analyzes the experiments of a simplified over-under combined nozzle and attempts to summarize the principles of the influence of twin-jet interaction on over-under combined nozzle performance. Firstly, twin-jet interaction directly changes the combined nozzle gross force via changing flow structure and parameters distribution inside the control volume. For example, the merging of supersonic jets forms wave systems, and the injection effect changes local pressure and the position of Free Shock Separation (FSS). Secondly, twin-jet interaction changes the force counting system. Assuming each flow-path as an isolated control volume leads to a mismatch of local pressure and reference ambient pressure. Thus, the combined nozzle should be considered as a whole control volume. Thirdly, twin-jet interaction may couple with jet-external interaction. On the one hand, jet-external interaction forms additional wave systems. On the other hand, the original expansion state of the nozzle during independent operation alters the performance baseline. Three practical experiments are conducted to verify these principles. These principles indicate that the influence of twin-jet interaction should be carefully considered in combined nozzle design. Utilizing the beneficial effect of twin-jet interaction can increase both the thrust and the lift, making the overall performance of the combined nozzle superior to the sum of two individual nozzles. However, twin-jet interaction may also decrease wall pressure, or alter the FSS position and flow direction of the over-expanded jet, resulting in a decrease in overall performance. These potential adverse effects should be avoided.

The conventional force measurement system using the balance has the problem of only being able to measure the overall load but unable to decouple the contribution of individual components, which limits the exploration of the flow mechanism behind aerodynamic load. With the development of Particle Image Velocimetry (PIV) technology，the PIV-based aerodynamic load measurement technology was developed, which can indirectly measure aerodynamic load by reconstructing multiple physical fields such as pressure and density. However, the performances of conventional pressure reconstruction methods were severely deteriorated when applied to supersonic flow fields. This drawback limits the application of reconstruction measurement methods in supersonic nozzles. To resolve this issue, a method for reconstructing the supersonic pressure field based on the Flux Vector Splitting (FVS) technique was proposed, and a thrust performance measurement scheme for supersonic nozzles based on PIV was established. PIV experiments on the Single Expansion Ramp Nozzle (SERN) were conducted using the direct-connect wind tunnel. Multiple physical fields and aerodynamic performance parameters such as flow rate, thrust, and lift were reconstructed under typical operating conditions. The evaluation results show that the data reconstructed by the FVS method possessed higher accuracy and better self-consistency, satisfying the laws of mass and momentum conservation. The relative errors of thrust and lift under overexpansion conditions were only −1.70% and 0.60%, respectively. The local errors of wall pressure after shock wave were lower than 3%. The performance of the proposed method is shown to be better than that of the conventional Poisson method and Spatial Integration (SI) method. Therefore, the experimental results verify the feasibility and high accuracy of the PIV-based thrust performance reconstruction measurement method when applied in supersonic nozzles, which can provide effective data supplementation to force measurement with the balance.

In response to the urgent requirements for wide-range high performance, geometric adaptability of the exit, and efficient thrust vectoring capability of axisymmetric nozzles for wide-range high-speed aircraft, an adjustable vector nozzle and an internal-middle-outer three-ring adjustment scheme were designed. Numerical simulations were conducted on the nozzle under typical operating conditions, and the flow field structure and performance of the nozzle under different modes were analyzed. The results show that by adjusting the positions of the middle and outer rings through translational motion, the nozzle can adjust the exit area and has the thrust vectoring capability. In the non-thrust vectoring mode, due to the adjustable exit area of the nozzle, the overexpansion phenomenon at low nozzle pressure ratio is effectively alleviated, and the thrust performance is significantly improved. Compared with the fixed geometry nozzle, the thrust coefficient can be increased by up to 32.75%. In the thrust vectoring mode, the maximum vector angle larger than 10° can be generated at Mach number 7, while the thrust coefficient is still greater than 0.92. To further reduce the base drag of the nozzle at low Mach number, a flow control method for secondary flow intake was proposed, which can reduce the base drag by up to 78.5% by reducing the internal and external pressure difference of the secondary flowpath. Finally, a wind tunnel scale-down experiment was conducted. The experiment results were in good agreement with the numerical simulation results, verifying the effectiveness of the proposed design scheme.

Bypass Dual Throat Nozzle (BDTN) is a novel type of fluidic thrust vectoring nozzle. To improve the infrared stealth performance of BDTN, a nozzle based on BDTN is proposed and numerically simulated. Each cross-section along the x-axis of the novel nozzle becomes a trapezoid, which is named “BDTN-TRA.” The main numerical simulation results show that BDTN-TRA can produce a thrust vectoring angle when the upper or lower bypass valve is open. The angle difference between the two conditions mentioned above is usually approximately 1°–2°. Even if the two bypasses are closed, BDTN-TRA can produce a small thrust vectoring angle at around 3°–5°. When the sidewall angle increases from 60° to 90°, the thrust coefficient and thrust vectoring angle under each work condition usually decrease. A larger aspect ratio indicates better performance. As the aspect ratio increases over 7.2, the performance of BDTN-TRA is quite close to that of BDTN with rectangular cross-sections at the same aspect ratio. These features will benefit the control and trimming for future aircraft design, especially for the flying wing layout aircraft. Last but not least, BDTN-TRA has a more extraordinary mixing performance compared with BDTN. The distributions of static temperature and axial velocity along the x-axis of BDTN-TRA with sidewall angle of 60° decrease faster than those of BDTN. When the total temperature of the inlet equals 1600 K, the static temperature difference between BDTN-TRA with sidewall angles of 60° and 90° is over 360 K at twice the length of the nozzle downstream of the nozzle exit, which is the reflection for excellent infrared stealth for the fighter.