Non metallic fuel settling valve is an aircraft fuel system equipment used to discharge water and fuel de-posited at the bottom of the aircraft fuel tank.The sedimentation valve is usually installed on the outer surface of the aircraft,and according to airworthiness regulations,it needs to meet fire resistance requirements.However,tradi-tional fire-resistant or fire-resistant designs(using high-temperature resistant metal materials,spraying fire-resistant coatings or topcoats,and using fire-resistant silicone rubber materials)are not suitable for sedimentation valves.To improve the fire resistance of non-metallic fuel settling valves,it is needed to start with the structure it-self.The transient thermal analysis of the valve with different body thicknesses is carried out.The influence of the thickness of the valve body on the combustion heat transfer and fuel seal is analyzed,and the results are verified by experiments.The results show that when the valve body is thickened to at least 13 mm,the valve begins to have fire resistance,that is,in the case of high temperature flame combustion,it can prevent either fuel leakage or over-temperature in the fuel tank.
In the process of air refueling,the oil joint of the tanker aircraft may break away from the air refueling hose and collide with the receiver aircraft,which can lead to significant damage to the receiver aircraft components and endanger the safety.In order to ensure the anti-impact ability of the receiver aircraft components,the dynamic simulation of the collision process between different parts of the receiver aircraft and the oil joint is carried out by using a transient nonlinear impact analysis,and the effects of different impact attitudes and different impact veloci-ties of the oil joint on the impact damage of the receiver aircraft components are studied.The results show that the vertical impact of the oil joint head is more destructive than that of the oil joint tail.When the oil joint head impacts the windshield or the upper wall panel vertically,the center of the leading edge is the weakest location for the verti-cal impact.It is also found that the stiffness matching between the structural components should be considered in the anti-impact design of the receiver aircraft.
It is necessary to study the influence of structural parameters of fuel bent pipes on the natural frequency of aircraft engines for the investigation of the vibration modal characteristics of aviation fuel bent pipes.Firstly,the modal analysis of a simple fixed support pipeline is conducted through simulation experiments.Secondly,the force hammer method is used for testing to verify the accuracy of the simulation data.Thirdly,the finite element mode-ling is conducted on fuel bent pipes with different external assembly parameters,and the changes in modal shapes and natural frequencies of each order are analyzed.Finally,based on the principles of dynamics,the reasons for the variation of natural frequency with assembly parameters are explained.The results show that the first mode of vibra-tion of the bent pipe is a bending deformation,with a maximum deformation of 96.63 mm,and the main vibration direction is in the Z direction;The second stage is the second bending deformation,with a maximum deformation of 78.20 mm,and the main vibration direction is in the Y direction.Radius,span,and bending radius are the main structural parameters that affect the natural frequency of bent pipes.
The coaxial rigid rotor configuration offers compact design,high hover efficiency,and high-speed poten-tial,but its coupled rotor-fuselage dynamics remain a key bottleneck.This paper reviews progress and challenges in three areas:theoretical modeling,experimental validation,and engineering application.For theoretical model-ing,the applicable boundaries and accuracy-efficiency trade-offs of momentum/vortex methods versus CFD/CSD coupling are reviewed,the aero-servo-elastic(ASE)effects at high speeds are clarified.For experimental valida-tion,the conflict between Froude and Mach similarity that leads to a lack of high-speed aeroelastic data is ana-lyzed,and non-contact measurements and hardware-in-the-loop(HIL)simulation are introduced.For engineering application,the heavy reliance of vibration reduction on empirical iteration is sort out,and digital twins for virtual sensing are explored.The review shows that high-fidelity simulations remain too costly,calling for nonlinear re-duced-order models(ROM)and multi-fidelity data fusion.The similitude conflict also hampers model validation.A future shift from experience-driven iteration to high-confidence model-driven design is needed,with an integrat-ed digital chain combining aero-servo-elastic modeling,hybrid(physical-virtual)validation,and analysis-driven certification(CbA).
Since ramp control is the most prevalent control method in complex helicopter maneuvering scenarios,in-vestigating rotor loads during control ramp increase holds significant guiding importance.By adopting free wake method and L-B dynamic stall model,combined with blade structural dynamics model,an aeroelastic coupling model applicable to rotor control ramp increase is established.The response of rotor aerodynamic and structural loads during control ramp increase in hovering and forward flight state of helicopter is calculated,and the effects of different initial controls and ramp increase rates on the rotor loads are analyzed.Results indicate that during the pro-cess of collective pitch increase in hover,increasing the initial collective pitch reduces the overshoot in thrust coeffi-cient and torsional moment,but amplifies the flap bending moment overshoot and the chord bending moment oscil-lations;Increasing the increase rate intensifies the overshoot and oscillations of rotor loads and accelerates the re-sponse of the chord bending moment.During longitudinal cyclic pitch ramp increase in hover,there is no significant overshoot in the pitching moment,but the oscillation amplitude of flap bending moment increases with the increase rate.During collective pitch increase in forward flight,increasing the increase rate accelerates the responses of the thrust coefficient and rolling moment and exacerbates the overshoot of the thrust coefficient and pitching moment.An increase in forward flight speed leads to more severe oscillations of the thrust coefficient,accelerates the re-sponse of the rolling moment,and enhances the overshoot phenomenon of the pitching moment.
