Conceptual Design Methodology For Near Space High Supersonic Unmanned Aerial Vehicle

The near-space High Supersonic Unmanned Aerial Vehicle(HSUAV)is a vehicle that can fly at high supersonic(3.0 < M < 5.0)speeds to perform specific tasks in near space.For such a new concept aircraft,a critical problem at the HSUAV conceptual design stage is how to determine the reasonable primary sizing parameters,including thrust loading,wing loading and maximum take-off weight of the HSUAV.The traditional aircraft sizing method is unable to solve this problem effectively.Therefore,this thesis intends to propose and investigate a new sizing methodology for the conceptual design of the HSUAV.The main works of this thesis are the following:1)Following the HSUAV design background,a request for proposal(RFP)of the HSUAV which includes the flight mission profiles and performance requirements is developed.On the basis of the RFP,the propulsion system type and aerodynamic configuration are determined by using the fundamentals of engine thermodynamic cycle and aerodynamics.The HSUAV features the turbine based combined cycle(TBCC)engine and normal blended wing fuselage with high swept wing.2)A modified design methodology using certain sizing parameters for aircraft sizing is proposed on the basis of the current constraint analysis and mission analysis methods.Following the methods,the confidence level of aircraft sizing results is enhanced by use of the wider-applicability and higher-accuracy aerodynamic and propulsion models.An iteration strategy for the HSUAV sizing is used.The traditional sizing methodology is used in the first iteration to obtain the initial design point.Then the aerodynamic and propulsion models with the wider-applicability and higher-accuracy are applied to improve confidence level of the results of HSUAV sizing.The modified aircraft sizing methodology expands the application range of the current constraint analysis and mission analysis methods.3)For the given propulsion system type in the HSUAV configuration,two analysis models for the tandem TBCC are established,including the engineering estimation model and thermodynamic cycle model.Based on these two analysis models,a MATLAB code is developed to estimate the thrust elapse coefficient and thrust specific fuel consumption of the TBCC in the flight mission profiles.In addition,the two-dimensional section of designed supersonic inlet of HSUAV is analyzed and optimized through using the isentropic gas flow theory.4)According to the RFP of HSUAV,the preliminary aerodynamic configuration design of HSUAV is completed.A convex leading edge design concept for nose-body is proposed to solve the issue of blended inlet body.An engineering estimation method is established for aerodynamic analysis models,as well as the numerical method.The numerical analysis tool consists of four parts,including the parametric geometry modelling code for aerodynamic configuration,three-dimensional unstructured grid generation code,flow solver and the code of the aerodynamic results analysis.The aspect ratio and dihedral angle of the HSUAV configuration are optimized by this numerical analysis tool,and the preferred aerodynamic configuration is finally obtained.5)The sizing process for the HSUAV conceptual design is completed by use of the modified sizing methodology,the established propulsion system model,and the preferred aerodynamic configuration.The design point of the TBCC is determined by qualitative analysis based on the RFP of HSUAV.Three iterations are used to complete the HSUAV sizing process.The confidence levels of aircraft sizing results are gradually increased with the iterations from the first round to the third round.In the end,the reasonability of the sizing results of the HSUAV is analyzed and validated.6)According to the sizing results of the HSUAV,the propulsion system model,and the preferred aerodynamic configuration,the conceptual design scheme of the HSUAV is presented,which includes the three-dimensional configuration CAD model,preliminary layout CAD model,weight features,aerodynamic characteristics,propulsion system characteristics and flight performance.The final results of the investigation in this thesis indicate that the proposed methodology can effectively solve the sizing problem in HSUAV conceptual design.This methodology can also be used for sizing other new concept aircraft.