Transactions of Nanjing University of Aeronautics & Astronautics
Issue 1,2026 Table of Contents
1 Investigation on Aerodynamic Interaction of Tandem Tilt-Wing and Multi-rotor in Transition Condition
2026(1):1-14.
DOI: 10.16356/j.1005-1120.2026.01.001
Abstract:
The complex aerodynamic interaction between tandem tilt-wing and multi-rotor directly affects the wing surface flow and rotor thrust， making it a critical factor during the tilt transition process of this configuration of rotorcraft. The aerodynamic interaction of tandem tilt-wing and multi-rotor is investigated based on the CFD method. The aerodynamic effect of multi tilt-rotor is simulated as virtual disk modeling by adding source terms to the Navier-Stokes equations， effectively reducing the calculation time while maintaining the accuracy of aerodynamic interaction calculations. Aerodynamic forces and flow field characteristics of the tandem tilt-wing and multi-rotor under different tilt angles are compared between cases with and without aerodynamic interaction. Furthermore， the differences in aerodynamic forces between dynamic tilt transition and fixed-angle conditions were compared. The results show that the aerodynamic interaction of multi-rotor obviously increases the lift of front tilt-wing at different tilt angles， the wing lift under interaction is increased by more than 40% compared with isolated wing at tilt angle of 15°for the computation in this paper， which is related to the increase of wing flow velocity and the suppression of flow separation caused by multi-rotor； the wing blocking effect will increase rotor thrust， especially near the tilt angles of 30° and 45°； the increases of rear wing lift and rear rotor thrust under aerodynamic interaction are not significant because of suppression by the front wing’s downwash； the unsteady effects during dynamic tilting have a relatively minor impact on aerodynamic interaction， with the aerodynamic forces on the rotors and wings during the dynamic tilting process showing little difference from those under corresponding fixed tilt angles.
The complex aerodynamic interaction between tandem tilt-wing and multi-rotor directly affects the wing surface flow and rotor thrust， making it a critical factor during the tilt transition process of this configuration of rotorcraft. The aerodynamic interaction of tandem tilt-wing and multi-rotor is investigated based on the CFD method. The aerodynamic effect of multi tilt-rotor is simulated as virtual disk modeling by adding source terms to the Navier-Stokes equations， effectively reducing the calculation time while maintaining the accuracy of aerodynamic interaction calculations. Aerodynamic forces and flow field characteristics of the tandem tilt-wing and multi-rotor under different tilt angles are compared between cases with and without aerodynamic interaction. Furthermore， the differences in aerodynamic forces between dynamic tilt transition and fixed-angle conditions were compared. The results show that the aerodynamic interaction of multi-rotor obviously increases the lift of front tilt-wing at different tilt angles， the wing lift under interaction is increased by more than 40% compared with isolated wing at tilt angle of 15°for the computation in this paper， which is related to the increase of wing flow velocity and the suppression of flow separation caused by multi-rotor； the wing blocking effect will increase rotor thrust， especially near the tilt angles of 30° and 45°； the increases of rear wing lift and rear rotor thrust under aerodynamic interaction are not significant because of suppression by the front wing’s downwash； the unsteady effects during dynamic tilting have a relatively minor impact on aerodynamic interaction， with the aerodynamic forces on the rotors and wings during the dynamic tilting process showing little difference from those under corresponding fixed tilt angles.
2026(1):15-26.
DOI: 10.16356/j.1005-1120.2026.01.002
Abstract:
A flow transition prediction method for calculating effects of Reynolds numbers on aerodynamic characteristics of airfoil is developed， and the accuracy of the method is verified by wind tunnel experiment data and other calculation results. On these basics， the infrared thermal imager experiment results of the flow transition in low-speed wind tunnel and the aerodynamic characteristics experiment results with variable Reynolds number in high-speed wind tunnel are carried out respectively， and compared with the numerical results of helicopter rotor airfoil. Specially， effects of Reynolds numbers on airfoil aerodynamic characteristics by means of flow transition under different working conditions are researched and some meaningful conclusions are obtained. The calculation method， experiment method and results as well as the flow transition analysis conclusions for aerodynamic characteristics can be used for the design of helicopter rotor airfoil， especially for the helicopters under the high altitude and low Reynolds number working conditions.
A flow transition prediction method for calculating effects of Reynolds numbers on aerodynamic characteristics of airfoil is developed， and the accuracy of the method is verified by wind tunnel experiment data and other calculation results. On these basics， the infrared thermal imager experiment results of the flow transition in low-speed wind tunnel and the aerodynamic characteristics experiment results with variable Reynolds number in high-speed wind tunnel are carried out respectively， and compared with the numerical results of helicopter rotor airfoil. Specially， effects of Reynolds numbers on airfoil aerodynamic characteristics by means of flow transition under different working conditions are researched and some meaningful conclusions are obtained. The calculation method， experiment method and results as well as the flow transition analysis conclusions for aerodynamic characteristics can be used for the design of helicopter rotor airfoil， especially for the helicopters under the high altitude and low Reynolds number working conditions.
2026(1):27-40.
DOI: 10.16356/j.1005-1120.2026.01.003
Abstract:
This study is dedicated to numerically investigate the dynamic behavior of a missile released from a helicopter under the influence of downwash from the rotating rotors using AGM-114 Hellfire and UH-60 as the testcase. Simulations are conducted using unsteady Reynolds-averaged Navier-Stokes （URANS） with shear stress transport （SST） k-ω turbulence model， incorporating six-degree-of-freedom （6-DOF） motion and overset grid. Two releasing scenarios， viz.， hover and forward flight， are analyzed under varying missile launch thrust and helicopter forward flight speed. Results reveal that the rotor downwash significantly affects the stability of the missile， particularly during hovering case， where low thrust prolongs wake interaction. In forward flight， the increased airspeed can in principle reduce wake influence but introduces asymmetrical aerodynamic effects on the trajectory of the missile. The findings offer guidance for missile release and launch planning in rotorcraft operations.
This study is dedicated to numerically investigate the dynamic behavior of a missile released from a helicopter under the influence of downwash from the rotating rotors using AGM-114 Hellfire and UH-60 as the testcase. Simulations are conducted using unsteady Reynolds-averaged Navier-Stokes （URANS） with shear stress transport （SST） k-ω turbulence model， incorporating six-degree-of-freedom （6-DOF） motion and overset grid. Two releasing scenarios， viz.， hover and forward flight， are analyzed under varying missile launch thrust and helicopter forward flight speed. Results reveal that the rotor downwash significantly affects the stability of the missile， particularly during hovering case， where low thrust prolongs wake interaction. In forward flight， the increased airspeed can in principle reduce wake influence but introduces asymmetrical aerodynamic effects on the trajectory of the missile. The findings offer guidance for missile release and launch planning in rotorcraft operations.
2026(1):40-54.
DOI: 10.16356/j.1005-1120.2026.01.004
Abstract:
Focusing on the unclear mechanism of aerodynamic interference in overlapping rotors of heavy-load electric vertical take-off and landing （eVTOL） aircraft， this paper aims to reveal the aerodynamic interference characteristics and flow field evolution laws of overlapping rotor configurations in hovering conditions through numerical simulation methods. The research method involves constructing a computational model for rotor flow fields and aerodynamic characteristics based on the Reynolds-averaged Navier-Stokes （RANS） equations and the Spalart-Allmaras （S-A） turbulence model. The dynamic simulation of rotor rotational motion was achieved by using the moving nested grid technology. The reliability of the computational method was ensured through the grid independence verification and the comparison with experimental data. The research results indicate that in overlapping rotor systems， rotor Ⅱ experiences a decrease in thrust， significant power fluctuations， and reduced hovering efficiency due to continuous interference from the adjacent rotor’s wake and blade-vortex interactions. Blade-tip vortices undergo breakage， fusion， and secondary rolling in the overlapping region， forming large-scale turbulent structures that lead to attenuation of the induced velocity field and aerodynamic efficiency losses. Additionally， the interaction between the rotor downwash and the fuselage triggers a “fountain effect” and a sudden increase in surface pressure on the fuselage， exacerbating flow field distortion. Based on the aforementioned mechanisms， the safe flight of overlapping rotor configurations can be achieved by optimizing the configuration strategy of the rotational speed phase difference between adjacent blades. This study provides a theoretical basis for the rotor layout design and the aerodynamic performance enhancement of heavy-load eVTOL aircraft.
Focusing on the unclear mechanism of aerodynamic interference in overlapping rotors of heavy-load electric vertical take-off and landing （eVTOL） aircraft， this paper aims to reveal the aerodynamic interference characteristics and flow field evolution laws of overlapping rotor configurations in hovering conditions through numerical simulation methods. The research method involves constructing a computational model for rotor flow fields and aerodynamic characteristics based on the Reynolds-averaged Navier-Stokes （RANS） equations and the Spalart-Allmaras （S-A） turbulence model. The dynamic simulation of rotor rotational motion was achieved by using the moving nested grid technology. The reliability of the computational method was ensured through the grid independence verification and the comparison with experimental data. The research results indicate that in overlapping rotor systems， rotor Ⅱ experiences a decrease in thrust， significant power fluctuations， and reduced hovering efficiency due to continuous interference from the adjacent rotor’s wake and blade-vortex interactions. Blade-tip vortices undergo breakage， fusion， and secondary rolling in the overlapping region， forming large-scale turbulent structures that lead to attenuation of the induced velocity field and aerodynamic efficiency losses. Additionally， the interaction between the rotor downwash and the fuselage triggers a “fountain effect” and a sudden increase in surface pressure on the fuselage， exacerbating flow field distortion. Based on the aforementioned mechanisms， the safe flight of overlapping rotor configurations can be achieved by optimizing the configuration strategy of the rotational speed phase difference between adjacent blades. This study provides a theoretical basis for the rotor layout design and the aerodynamic performance enhancement of heavy-load eVTOL aircraft.
5 A Dynamic Matrix Controller with Feedforward for Flow Field in Intermittent Transonic Wind Tunnels
2026(1):55-72.
DOI: 10.16356/j.1005-1120.2026.01.005
Abstract:
Intermittent transonic wind tunnels demand high-precision and stable control of Mach number and stagnation pressure during the variation of model angle of attack， while the traditional proportional integral derivative （PID） control strategy fails to achieve the Mach number control error target of 0.001 and is inept at resisting flow field disturbances caused by rapid changes in angle of attack. Aiming at this problem， this study takes the intermittent transonic wind tunnel of China Aerodynamics Research and Development Center as the research object and optimizes the wind tunnel control system for its characteristics of multi-input multi-output （MIMO）， large time delay and nonlinearity. First， the control system structure is reconstructed by introducing ejection pressure as a controlled variable to reduce the time lag from the main pressure regulating valve to the test section， and selecting static pressure instead of Mach number as a controlled variable to weaken the nonlinear coupling between Mach number and static pressure. Second， based on the first-order plus dead time process model of the wind tunnel， a MIMO dynamic matrix controller （DMC） for wind tunnel flow field is designed. Considering the predictable nature of angle-of-attack changes， a feedforward compensation strategy is integrated into the DMC to mitigate the pressure disturbance in the test section caused by the variation of angle of attack. Third， a complete tuning strategy for DMC parameters including sampling time， prediction horizon， control horizon and weight coefficients is formulated according to the wind tunnel model parameters under different Mach numbers. Experimental validations through practical blowing tests are carried out at Mach numbers of 0.578， 0.675， 0.714 and 0.822 to compare the control performance of the proposed feedforward DMC strategy with the traditional PID control and DMC without feedforward compensation. The results show that the proposed strategy stably controls the Mach number error within 0.001， and significantly improves the stagnation pressure control accuracy and anti-disturbance capability of the wind tunnel flow field. Moreover， the strategy exhibits excellent repeatability and robustness under different Mach number conditions. This study effectively solves the problem of high-precision flow field control under the rapid change of angle of attack in intermittent transonic wind tunnels， and provides a technical reference for the flow field control of complex fluid test devices such as hypersonic wind tunnels.
Intermittent transonic wind tunnels demand high-precision and stable control of Mach number and stagnation pressure during the variation of model angle of attack， while the traditional proportional integral derivative （PID） control strategy fails to achieve the Mach number control error target of 0.001 and is inept at resisting flow field disturbances caused by rapid changes in angle of attack. Aiming at this problem， this study takes the intermittent transonic wind tunnel of China Aerodynamics Research and Development Center as the research object and optimizes the wind tunnel control system for its characteristics of multi-input multi-output （MIMO）， large time delay and nonlinearity. First， the control system structure is reconstructed by introducing ejection pressure as a controlled variable to reduce the time lag from the main pressure regulating valve to the test section， and selecting static pressure instead of Mach number as a controlled variable to weaken the nonlinear coupling between Mach number and static pressure. Second， based on the first-order plus dead time process model of the wind tunnel， a MIMO dynamic matrix controller （DMC） for wind tunnel flow field is designed. Considering the predictable nature of angle-of-attack changes， a feedforward compensation strategy is integrated into the DMC to mitigate the pressure disturbance in the test section caused by the variation of angle of attack. Third， a complete tuning strategy for DMC parameters including sampling time， prediction horizon， control horizon and weight coefficients is formulated according to the wind tunnel model parameters under different Mach numbers. Experimental validations through practical blowing tests are carried out at Mach numbers of 0.578， 0.675， 0.714 and 0.822 to compare the control performance of the proposed feedforward DMC strategy with the traditional PID control and DMC without feedforward compensation. The results show that the proposed strategy stably controls the Mach number error within 0.001， and significantly improves the stagnation pressure control accuracy and anti-disturbance capability of the wind tunnel flow field. Moreover， the strategy exhibits excellent repeatability and robustness under different Mach number conditions. This study effectively solves the problem of high-precision flow field control under the rapid change of angle of attack in intermittent transonic wind tunnels， and provides a technical reference for the flow field control of complex fluid test devices such as hypersonic wind tunnels.
2026(1):73-85.
DOI: 10.16356/j.1005-1120.2026.01.006
Abstract:
Focusing on civil aircraft flap skew detection design， this paper proposes a high-robustness monitoring design methodology to address insufficient monitor robustness that may trigger false alarms and disrupt airline operations. Based on flap skew detection principles and threshold design criteria， the threshold range is defined with upper limit of maximum deformation under aerodynamic load and lower limit of sensor error margin and nominal flight deformation. Since the complex loading conditions of maximum flap differential deformation （max Δλ） during normal flight cannot be theoretically determined， probabilistic methods are employed： Flight test data from hundreds of sorties are analyzed using generalized extreme value distribution. Confidence levels are verified via Kolmogorov-Smirnov （K-S） hypothesis testing. Then probability density function of max Δλ is established. The false alarm rate is calculated through cumulative probability values of max Δλ at varying thresholds. Boundary conditions for false alarm rate are determined by safety assessment and dispatch reliability analysis. The derived monitoring threshold is verified against finite element analysis predictions and iron bird rig test. The results confirm the methodology’s validity， meeting all design objectives.
Focusing on civil aircraft flap skew detection design， this paper proposes a high-robustness monitoring design methodology to address insufficient monitor robustness that may trigger false alarms and disrupt airline operations. Based on flap skew detection principles and threshold design criteria， the threshold range is defined with upper limit of maximum deformation under aerodynamic load and lower limit of sensor error margin and nominal flight deformation. Since the complex loading conditions of maximum flap differential deformation （max Δλ） during normal flight cannot be theoretically determined， probabilistic methods are employed： Flight test data from hundreds of sorties are analyzed using generalized extreme value distribution. Confidence levels are verified via Kolmogorov-Smirnov （K-S） hypothesis testing. Then probability density function of max Δλ is established. The false alarm rate is calculated through cumulative probability values of max Δλ at varying thresholds. Boundary conditions for false alarm rate are determined by safety assessment and dispatch reliability analysis. The derived monitoring threshold is verified against finite element analysis predictions and iron bird rig test. The results confirm the methodology’s validity， meeting all design objectives.
2026(1):85-94.
DOI: 10.16356/j.1005-1120.2026.01.007
Abstract:
In fatigue damage tolerance verification tests of aircraft structures， the simulation and loading of flight-by-flight spectra require considerable time and resources. To improve the efficiency of load spectrum design and testing， an equivalent constant-amplitude spectrum design method for flight-by-flight spectra is proposed based on the equivalence of crack growth behavior. By combining the Paris crack growth model with the Walker stress ratio correction， the equivalent stress amplitude is directly calculated using structural parameters and load spectrum characteristics， enabling a rapid transformation from variable-amplitude spectra to constant-amplitude spectra. The original spectrum is discretized based on the load-exceedance curve， and the equivalence relationship between multi-level block spectra and constant-amplitude spectra is established. Taking a typical lower wing skin structure of a transport aircraft as an example， two equivalent spectra are designed and validated through fatigue crack growth tests on 2024-T351 center-hole plate specimens. The experimental results show that the fatigue life deviation between the equivalent spectra and the original flight-by-flight spectrum is within 10%， demonstrating the effectiveness of the proposed method. Moreover， the equivalent spectrum constructed under the condition of invariant mean flight stress exhibits higher equivalence accuracy. The influence of spectral shape on the equivalent stress amplitude is further analyzed， revealing that the equivalent stress amplitude increases with the spectrum shape coefficient. The proposed method provides a useful reference for load spectrum design in aircraft structural damage tolerance verification tests.
In fatigue damage tolerance verification tests of aircraft structures， the simulation and loading of flight-by-flight spectra require considerable time and resources. To improve the efficiency of load spectrum design and testing， an equivalent constant-amplitude spectrum design method for flight-by-flight spectra is proposed based on the equivalence of crack growth behavior. By combining the Paris crack growth model with the Walker stress ratio correction， the equivalent stress amplitude is directly calculated using structural parameters and load spectrum characteristics， enabling a rapid transformation from variable-amplitude spectra to constant-amplitude spectra. The original spectrum is discretized based on the load-exceedance curve， and the equivalence relationship between multi-level block spectra and constant-amplitude spectra is established. Taking a typical lower wing skin structure of a transport aircraft as an example， two equivalent spectra are designed and validated through fatigue crack growth tests on 2024-T351 center-hole plate specimens. The experimental results show that the fatigue life deviation between the equivalent spectra and the original flight-by-flight spectrum is within 10%， demonstrating the effectiveness of the proposed method. Moreover， the equivalent spectrum constructed under the condition of invariant mean flight stress exhibits higher equivalence accuracy. The influence of spectral shape on the equivalent stress amplitude is further analyzed， revealing that the equivalent stress amplitude increases with the spectrum shape coefficient. The proposed method provides a useful reference for load spectrum design in aircraft structural damage tolerance verification tests.
2026(1):95-109.
DOI: 10.16356/j.1005-1120.2026.01.008
Abstract:
The buoyancy-induced flow constitutes a core scientific issue for thermal management of electronic devices and thermal design of energy systems， where accurate characterization of flow and heat transfer is essential to improve thermal efficiency. In this work， buoyancy-induced flow above two heating elements flush-mounted at the bottom of a square enclosure containing air is numerically investigated over a range of Rayleigh numbers （0 < R a ≤ 1.5 × 10 8 ![]()
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The buoyancy-induced flow constitutes a core scientific issue for thermal management of electronic devices and thermal design of energy systems， where accurate characterization of flow and heat transfer is essential to improve thermal efficiency. In this work， buoyancy-induced flow above two heating elements flush-mounted at the bottom of a square enclosure containing air is numerically investigated over a range of Rayleigh numbers （
2026(1):110-126.
DOI: 10.16356/j.1005-1120.2026.01.009
Abstract:
With the rapid development of the aviation industry， air travel has become one of the most important modes. Improving the service quality of civil aviation airports is crucial to their competitiveness. This study intends to develop a scientific and rational evaluation methodology and framework for assessing service quality in civil aviation airports， thereby providing a theoretical foundation and practical guidance for enhancing service standards in the aviation industry. First， the study constructs a CRITIC-bidirectional grey possibility clustering model， which uses the CRITIC method to determine the weights of indicators and integrates the forward grey possibility clustering model and the inverse grey possibility clustering model to determine possibility functions from two perspectives. Second， a service quality evaluation index system for civil airports is constructed from four dimensions， and the weights of each index within the system are subsequently calculated. Finally， the constructed model is applied to evaluate the service quality of nine domestic civil airports. Based on the clustering results， targeted countermeasures and suggestions are proposed. Empirical results demonstrate that， compared to the traditional grey possibility clustering model， the proposed model balances the objectivity of indicator weighting， the objectivity of possibility function construction， and the simplicity of the computational process， thereby possessing significant theoretical and practical implications.
With the rapid development of the aviation industry， air travel has become one of the most important modes. Improving the service quality of civil aviation airports is crucial to their competitiveness. This study intends to develop a scientific and rational evaluation methodology and framework for assessing service quality in civil aviation airports， thereby providing a theoretical foundation and practical guidance for enhancing service standards in the aviation industry. First， the study constructs a CRITIC-bidirectional grey possibility clustering model， which uses the CRITIC method to determine the weights of indicators and integrates the forward grey possibility clustering model and the inverse grey possibility clustering model to determine possibility functions from two perspectives. Second， a service quality evaluation index system for civil airports is constructed from four dimensions， and the weights of each index within the system are subsequently calculated. Finally， the constructed model is applied to evaluate the service quality of nine domestic civil airports. Based on the clustering results， targeted countermeasures and suggestions are proposed. Empirical results demonstrate that， compared to the traditional grey possibility clustering model， the proposed model balances the objectivity of indicator weighting， the objectivity of possibility function construction， and the simplicity of the computational process， thereby possessing significant theoretical and practical implications.
2026(1):127-144.
DOI: 10.16356/j.1005-1120.2026.01.010
Abstract:
Flight behavior analysis provides the fundamental basis for the future development of air traffic management （ATM）. The characteristics of aircraft behavior are inherently reflected in their flight trajectories， impacting flight efficiency and safety levels. However， existing research largely addresses inefficient or abnormal trajectories from a single perspective， with an absence of a unified evaluation standard. This paper introduces a method for analyzing flight deviation behavior based on automatic dependent surveillance-broadcast （ADS-B） data， defining novel metrics of trajectory redundancy and trajectory deviation. An adaptive detection algorithm is developed to capture diverse deviation patterns. Results reveal that higher trajectory redundancy is linked to lower operational efficiency， while trajectory deviation effectively identify stepped descents， holding patterns， detours， and other behaviors. The approach offers data-driven support for anomaly detection， performance evaluation and air traffic management， with substantial significance for civil aviation applications.
Flight behavior analysis provides the fundamental basis for the future development of air traffic management （ATM）. The characteristics of aircraft behavior are inherently reflected in their flight trajectories， impacting flight efficiency and safety levels. However， existing research largely addresses inefficient or abnormal trajectories from a single perspective， with an absence of a unified evaluation standard. This paper introduces a method for analyzing flight deviation behavior based on automatic dependent surveillance-broadcast （ADS-B） data， defining novel metrics of trajectory redundancy and trajectory deviation. An adaptive detection algorithm is developed to capture diverse deviation patterns. Results reveal that higher trajectory redundancy is linked to lower operational efficiency， while trajectory deviation effectively identify stepped descents， holding patterns， detours， and other behaviors. The approach offers data-driven support for anomaly detection， performance evaluation and air traffic management， with substantial significance for civil aviation applications.
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