Quantitative analysis of the morphing wing mechanism of raptors:IMMU-based motion capture system and its application on gestures of a Falco peregrinus

This paper presented a novel tinny motion capture system for measuring bird posture based on inertial and magnetic measurement units that are made up of micromachined gyroscopes, accelerometers, and magnetometers. Multiple quaternion-based extended Kalman filters were implemented to estimate the absolute orientations to achieve high accuracy.Under the guidance of ornithology experts, the extending/contracting motions and flapping cycles were recorded using the developed motion capture system, and the orientation of each bone was also analyzed. The captured flapping gesture of the Falco peregrinus is crucial to the motion database of raptors as well as the bionic design.

Time-Domain Analysis of Body Freedom Flutter Based on 6DOF Equation

The reduced weight and improved efficiency of modern aeronautical structures result in a decreasing separation of frequency ranges of rigid and elastic modes.Particularly,a high-aspect-ratio flexible flying wing is prone to body freedomflutter(BFF),which is a result of coupling of the rigid body short-periodmodewith 1st wing bendingmode.Accurate prediction of the BFF characteristics is helpful to reflect the attitude changes of the vehicle intuitively and design the active flutter suppression control law.Instead of using the rigid body mode,this work simulates the rigid bodymotion of the model by using the six-degree-of-freedom(6DOF)equation.A dynamicmesh generation strategy particularly suitable for BFF simulation of free flying aircraft is developed.An accurate Computational Fluid Dynamics/Computational Structural Dynamics/six-degree-of-freedom equation(CFD/CSD/6DOF)-based BFF prediction method is proposed.Firstly,the time-domain CFD/CSD method is used to calculate the static equilibrium state of the model.Based on this state,the CFD/CSD/6DOF equation is solved in time domain to evaluate the structural response of themodel.Then combinedwith the variable stiffnessmethod,the critical flutter point of the model is obtained.This method is applied to the BFF calculation of a flyingwing model.The calculation results of the BFF characteristics of the model agree well with those fromthe modalmethod andNastran software.Finally,the method is used to analyze the influence factors of BFF.The analysis results show that the flutter speed can be improved by either releasing plunge constraint or moving the center ofmass forward or increasing the pitch inertia.

A combined airfoil with secondary feather inspired by the golden eagle and its influences on the aerodynamics

Bird flight is a remarkable adaption that has allowed thousands of species to colonize all terrestrial habitats. A golden eagle has impressive flying abilities, such as hovering, perching, preying and attacking. To reveal the flying abilities, avian geometry of a golden eagle was extracted based on noncontact surface measurements using a ROMBER three-dimensional laser scanner. Distributions of a camber line, thickness and a secondary feather line of the extracted point cloud were fitted using convenient analytical expressions. A traditional airfoil was established with the camber line and thickness, then a combined airfoil was constructed by combining the traditional airfoil with a secondary feather. Oscillations of an airfoil as well as rapid pitch up were simplified as a sine wave around the quarter chord axis. Thereafter, both steady and unsteady aerodynamic performances of the airfoil are computed, the influences of the secondary feather on the steady and unsteady aerodynamics were further studied.

Leakage of an eagle flight feather and its influence on the aerodynamics

We investigate how the barb of bird feathers is changed along both the rachis and barb.To investigate the microstructures and the mechanical behaviors of barbs,a series of barbs are manually cut from an eagle’s primary feather to observe the cross sections.Aλ-like cross section with a tiny hook is observed at the right feet at each section.Afterwards,a measurement of the setup system is developed to evaluate the leakage ratio of a feather followed by a numerical predicting approach based on the CFD method.It is found that the air leakage increases linearly against the pressure,and the predicted results coincide well with the experimental results.Finally,the influences of leakage of the flight feather on both steady and unsteady aerodynamics are studied.

Quantitative analysis of the morphing wing mechanism of raptors:Morphing kinematics of Falco peregrinus wing

Raptors are getting more attention from researchers because of their excellent flight abilities.And the excellent wing morphing ability is critical for raptors to achieve high maneuvering flight,which can be a good bionic inspiration for unmanned aerial vehicles(UAV)design.However,morphing wing motions of Falco peregrinus with multi postures cannot be consulted since such a motion database was nonexistent.This study aimed to provide data reference for future research in wing morphing kinetics.We used the computed tomography(CT)approach to obtain nine critical postures of the Falco peregrinus wing skeleton,followed with motion analysis of each joint and bone.Based on the obtained motion database,a six-bar kinematic model was proposed to regenerate wing motions with a high fidelity.

Analysis on Physical Mechanism of Sound Generation inside Cavities Based on Acoustic Analogy Method

Analysis of coupling aerodynamics and acoustics are performed to investigate the self-sustained oscillation and aerodynamic noise in two-dimensional flow past a cavity with length to depth ratio of 2 at subsonic speeds. The large eddy simulation (LES) equations and integral formulation of Ffowcs-Williams and Hawings (FW-H) are solved for the cavity with same conditions as experiments. The obtained density-field agrees well with Krishnamurty’s experimental schlieren photograph, which simulates flow-field distributions and the direction of sound wave radiation. The simulated self-sustained oscillation modes inside the cavity agree with Rossiter’s and Heller’s predicated results, which indicate frequency characteristics are obtained. Moreover, the results indicate that the feedback mechanism that new shedding-vortexes induced by propagation of sound wave created by the impingement of the shedding-vortexes in the shear-layer and rear cavity face leads to self-sustained oscillation and high noise inside the cavity. The peak acoustic pressure occurs in the first oscillation mode and the most of sound energy focuses on the low-frequency region.

Shape reconstructions and morphing kinematics of an eagle during perching manoeuvres

The key to high manoeuvre ability in bird flight lies in the combined morphing of wings and tail.The perching of a wild Haliaeetus Albicilla without running or wing flapping is recorded and investigated using a high-speed digital video.A shape reconstruction method is proposed to describe wing contours and tail contours during perching.The avian airfoil geometries of the Aquila Chrysaetos are extracted from noncontact surface measurements using a ROMBER 3D laser scanner.The wing planform,chord distribution and twist distribution are fitted in convenient analytical expressions to obtain a 3D wing geometry.A three-jointed arm model is proposed to associate with the 3D wing geometry,while a one-joint arm model is proposed to describe the kinematics of tail.Therefore,a 3D bird model is established.The perching sequences of the wild eagle are recaptured and regenerated with the proposed 3D bird model.A quasi-steady aerodynamic model is applied in the aerodynamic predictions,a four-step Adams-Bashforth method is used to calculate the ordinary differential equations,thus a BFGS based optimization method is established to predict the perching motions.

Measuring the flexibility matrix of an eagle’s flight feather and a method to estimate the stiffness distribution

Flight feathers stand out with extraordinary mechanical properties for flight because they are lightweight but stiff enough.Their elasticity has great effects on the aerodynamics, resulting in aeroelasticity.Our primary task is to figure out the stiffness distribution of the feather to study the aeroelastic effects.The feather shaft is simplified as a beam, and the flexibility matrix of an eagle flight feather is tested.A numerical method is proposed to estimate the stiffness distributions along the shaft length based on an optimal Broyden–Fletcher–Goldfarb–Shanno(BFGS) method with global convergence.An analysis of the compressive behavior of the shaft based on the beam model shows a good fit with experimental results.The stiffness distribution of the shaft is finally presented using a 5 th order polynomial.

Quantitative analysis of the morphing wing mechanism of raptors: Analysis methods, folding motions, and bionic design of Falco Peregrinus

Raptors can change the shape and area of their wings to an exceptional degree in a fast and efficient manner,surpassing other birds,insects,or bats.Some researchers have focused on the functional properties of muscle skeletons,mechanics,and flapping robot design.However,the wing motion of the birds of prey has not been measured quantitatively,and synthetic bionic wings with morphing abilities similar to raptors are far from reality.Therefore,in the current study,a 3D suspension system for holding bird carcasses was designed and fabricated to fasten the wings of Falco Peregrinus with a series of morphing postures.Subsequently,the wing skeleton of the falcon was scanned during extending motions using the computed tomography(CT)approach to obtain three consecutive poses.Subsequently,the skeleton was reconstructed to identify the contribution of the forelimb bones to the extending/folding motions.Inspired by these findings,we propose a simple mechanical model with four bones to form a wing-morphing mechanism using the proposed pose optimisation method.Finally,a bionic wing mechanism was implemented to imitate the motion of the falcon wing—divided into inner and outer wings with folding and twisting motions.The results show that the proposed four-bar mechanism can track bone motion paths with high fidelity.

Experimental and computational investigation of transonic speed

The influence of the wing-tip vortex of leading aircraft on energy savings,quantified by formation aerodynamic force fraction of the following aircraft,is studied at transonic speed for a matrix of leading aircraft’s vortex locations.The research model adopts the hybrid formation of medium and large aircraft.The leading aircraft is scaled by 2.1%,and the following aircraft is scaled by 1.4%.An aerodynamic benefit "map"is developed to determine the optimum location of the following aircraft relative to the leading aircraft wake and to compare with experimental results,thus validating the use of CFD for the formation flight at cruising speed.The response surface model of aerodynamic gain effect relative to formation parameters is established via numerical calculation and wind tunnel test.The optimal formation parameters and the setting criteria of the study model are optimized.Results show that the wing-tip vortex of large aircraft significantly increases lift and reduces drag on the medium-sized aircraft following it.Reduced drag slightly increases with the flow direction position.With the increase of flow direction distance,the peak area moves from 15%of wing-tip overlap to 20%of overlap.In addition,the maximum drag decreases about 16%,and the maximum lift increases about 12%.The lift drag ratio of the optimal position is increased by 27%,which is twice as large as that of the same scale ratio aircraft formation.Results show that the increase of lift is mainly caused by the increase of suction peak and suction range.

Dynamics,stability,and control of a four-cable mount system for wind tunnel test

A four-cable mount system is proposed for full-model wind tunnel flutter tests,which may adjust the pitch and roll attitude of the aircraft scaled model and ensure that the model is not subjected to cable tension.The system provides sufficient support to simulate the free flight of the aircraft by applying appropriate spring stiffness and cable tensions.The proposed fourcable mount system is modeled based on Lagrange mechanics,and its dynamics equations consider aerodynamic effects.The singularity of the system and its bifurcation characteristics under flow conditions are analysed to determine the supercritical bifurcation phenomenon for different tension levels and distances from the front suspension point to the mass centre of the model.The mathematical expressions of the longitudinal flight stability of the cable mount system are derived by linearising the system dynamics equations using small perturbations.The influence of the cable tension,spring stiffness,suspension point position,and other factors on the flight stability of the aircraft are analysed.A feedforward control algorithm is proposed to minimize the total elastic potential energy of the system.The results show that the model is in the level flight state when the elastic potential energy of the four-cable mount system is minimized.A feedback control design method is proposed based on the Lyapunov stability theory to derive the closed-loop stability conditions.The system dynamics model that includes the aircraft rigid body model,flexible cables,pulleys,springs,aerodynamic model,and servo motor control is established using the flexible multibody dynamics method.A multibody dynamics solver and Simulink are used to simulate the attitude adjustment of the model in the wind tunnel and verify the supercritical bifurcation characteristics of the system and the effectiveness of the feedback and feedforward control.

A modified-Liutex-based vortex-core-line extractor and its application

Vortices,which appear as swirling behaviour of a flow field,are an important phenomenon in fluid dynamics,and the extraction of vortex cores is necessary for the research of vortex evolution in many areas of fluid mechanics.The Liutex method is a milestone in vortex identification and provides a reasonable mathematical definition for a vortex core.Based on this definition,a novel integration-based method is presented,which can reduce the numerical error in the integration process through location optimization.Two typical test cases,the wake vortices of an A320 in the near-ground stage and a helicopter rotor,are examined to show that the proposed method can extract continuous vortex core lines with accuracy and efficiency for vortex parameter study.

Microjet flow control in an ultra-compact serpentine inlet

Microjets are used to control the internal flow to improve the performance of an ultra-compact serpentine inlet. A highly offset serpentine inlet with length-to-diameter ratio of 2.5 is designed and static tests are conducted to analyze the internal flow characteristics in terms of pressure recovery, distortion and flow separation. Flow separation is encountered in the second S-turn, and two strong counter-rotating vortices are formed at the aerodynamic interthce plane （AIP） face which occupy a quarter of the outlet area and result in severe pressure loss and distortion. A flow control model employing a row of microjets in the second turn is designed based on the internal flow characteristics and simplified CFD simulations. Flow control tests are conducted to verify the control effectiveness and understand the characteristics as a function of inlet throat Mach number, injection mass flow ratio, jet Mach number and momentum coefficient. At all test Mach numbers, microjet flow control （MFC） effectively improves the recovery and reduces the distortion intensity. Between inlet throat Mach number 0.2 and 0.5, the strong flow separation in the second S-turn is suppressed at an optimum jet flow ratio of less than 0.65%, resulting in a maximum improvement of 4% for pressure recovery coefficient and a maximum decrease of 75% for circumferential distortion intensity at cruise. However, in order to suppress the flow separation, the injection rate should retain in an effective range. When the injection rate is higher than this range, the flow is degraded and the distortion contour is changed from 90° circumferential distortion pattern to 180° circumferential distortion pattern. Detailed data analysis shows that this optimum flow ratio depends on inlet throat Mach number and the monlcntunl coefficient affects the control effectiveness in a dual stepping manner.

Transonic wing stall of a blended flying wing common research model based on DDES method

Numerical simulation of wing stall of a blended flying wing configuration at transonic speed was conducted using both delayed detached eddy simulation（DDES） and unsteady Reynolds-averaged Navier-Stokes（URANS） equations methods based on the shear stress transport（SST） turbulence model for a free-stream Mach number 0.9 and a Reynolds number 9.6 × 10. A joint time step/grid density study is performed based on power spectrum density（PSD） analysis of the frequency content of forces or moments, and medium mesh and the normalized time scale0.010 were suggested for this simulation. The simulation results show that the DDES methods perform more precisely than the URANS method and the aerodynamic coefficient results from DDES method compare very well with the experiment data. The angle of attack of nonlinear vortex lift and abrupt wing stall of DDES results compare well with the experimental data. The flow structure of the DDES computation shows that the wing stall is caused mainly by the leeward vortex breakdown which occurred at x/x= 0.6 at angle of attack of 14°. The DDES methods show advantage in the simulation problem with separation flow. The computed result shows that a shock/vortex interaction is responsible for the wing stall caused by the vortex breakdown. The balance of the vortex strength and axial flow, and the shock strength, is examined to provide an explanation of the sensitivity of the breakdown location. Wing body thickness has a great influence on shock and shock/vortex interactions, which can make a significant difference to the vortex breakdown behavior and stall characteristic of the blended flying wing configuration.

Investigation of rotor–stator interaction broadband noise using a RANS-informed analytical method

A Reynolds-Averaged Navier Stokes(RANS)-information analytical method for predicting Rotor-Stator Interaction(RSI)broadband noise is established in this paper.First,the turbulence information is deduced from RANS simulation result.Then,the unsteady load on the stator blade is calculated using a strip theory approach based on LINearized SUBsonic unsteady flow in cascade(LINSUB)and 2-D equivalence method.In the end,the sound power of RSI broadband noise is calculated by coupling the unsteady load on the stator blade with acoustic analogy and annular duct mode.The broadband noise model part of the RANS-information analytical method is validated against the upstream sound power of an annular cascade experimental bench.Besides,the RANS-information analytical method is used in predicting RSI broadband noise of a single-stage axial fan acoustic experimental bench,the results illustrate that the RANS-information analytical method can accurately predict the RSI broadband noise in different fan working conditions.After simplification the Wave Leading Edge(WLE)stator blade,the effect of WLE stator blade on RSI broadband noise is studies.Although the simplification may bring some discrepancies,the results illustrate that the RANS-information analytical method has the capability for further studies on the broadband noise reduction with WLE stator blade.

Constrained layer damping treatment of a model support sting

In transonic wind tunnel tests,the pulsating airflow is prone to induce the first order resonance of the sting support system.The resonance limits the wind tunnel test envelope,makes the test data inaccurate,and brings potential security risks.In this paper,a model support sting with constrained layer damping(CLD)treatment is proposed to reduce the first order resonance response.The CLD treatment mainly consists of material selection and geometric optimization processes.The damping performance of the optimized CLD sting is compared with an AISI 1045 steel sting with the identical diameter in laboratory.The frequency response curves of the CLD sting support system and the AISI 1045 steel sting support system are obtained by sine sweep tests.The test results show that the first order resonance response of the CLD sting support system is 37.3%of that of the AISI 1045 steel sting support system.The first order damping ratios are calculated from the frequency response curves by half power point method.It is found that the first order damping ratio of the CLD sting support system is approximately 2.6 times that of the AISI 1045 steel sting support system.

Closed-loop flow control of an ultra-compact serpentine inlet based on nondimensional model

Ultra-compact serpentine inlet faces serve inlet-engine compatibility issues due to flow distortion.To ensure inlet-engine compatibility over a wide range of Mach number,novel active flow control techniques with the ability of being opened or adjusted as needed draw many attentions in recent years.In this paper,a feedback control system was developed based on the method of microjet blowing.The proposed system includes a pressure adjusting valve to adjust the control effort,a dynamic pressure sensor to sense the inlet distortion intensity,a signal processing instrument to calculate the Root-Mean-Squared(RMS)pressure,and a controller to implement feedback control.To achieve high quality closed-loop controls at dynamic conditions,a novel nondimensional feedback method was developed.The advantage of this nondimensional method was validated at both off-design and arbitrarily changing Mach number conditions.With a sectional PI control law,the RMS control error reduced more than 56%at arbitrary changing conditions.Works in this paper also showed that the dynamics of this nondimensional system can be simplified as a stable second-order overdamped system.

Utilization of Whale-inspired Leading-edge Tubercles for Airfoil Noise Reduction

Numerical studies are conducted to explore the noise reduction effect of leading-edge tubercles inspired by humpback whale flippers.Large eddy simulations are performed to solve the flow field,while the acoustic analogy theory is used for noise prediction.In this paper,a baseline airfoil with a straight leading-edge and three bionic airfoils with tubercled leading-edges are simulated.The tubercles have sinusoidal profiles and the profiles are determined by the tubercle wavelength and amplitude.The tubercles used in this study have a fixed wavelength of 0.1c with three different amplitudes of 0.1c,0.15c and 0.2c,where c is the mean chord of the airfoil.The freestream velocity is set to 40 m/s and the chord based Reynolds number is 400,000.The predicted flow field and acoustic field of the baseline airfoil are compared against the experiments and good agreements are found.A considerable noise reduction level is achieved by the leading-edge tubercles and the tubercle with larger amplitude can obtain better noise reduction.The underlying flow mechanisms responsible for the noise reduction are analyzed in detail.