Pitching Moment Research Articles

Investigation of Aerodynamic Performance of Multi-element Airfoil in Free Surface Effect

Abstract This study examines the aerodynamic performance of the 30P30N multi-element airfoil flying on a free surface, a numerical computational model is established, the governing equations are addressed through the application of the finite volume method, while the volume of fluid method is utilized for the simulation of two-phase flow. Different mesh strategies and turbulence models are compared, with the numerical methods validated against experimental data. The aerodynamic properties of the 30P30N multi-element airfoil on the free surface at different wavelengths and wave heights are given by numerical computations, which show that the ground effect of the 30P30N multi-element airfoil differ from those on a clean airfoil, the aerodynamic coefficients of the 30P30N multi-element airfoil decrease as flight altitude lowers, with the pitching and drag moment of the slat is opposite to that of the main airfoil and the flap. Additionally, when the 30P30N multi-element airfoil flies over a regular wave surface, the aerodynamic coefficients of each component exhibit periodic behavior.

Optimizing the Landing Stability of Blended-Wing-Body Aircraft with Distributed Electric Boundary-Layer Ingestion Propulsors through a Novel Thrust Control Configuration

The imperative for energy conservation and environmental protection has led to the development of innovative aircraft designs. This study explored a novel thrust control configuration for blended-wing-body (BWB) aircraft with distributed electric boundary-layer ingestion (BLI) propulsors, addressing the issues of sagging and altitude loss during landing. The research focused on a small-scale BWB demonstrator equipped with six BLI fans, each with a 90 mm diameter. Various thrust control configurations were evaluated to achieve significant thrust reduction while maintaining lift, including dual-layer sleeve, separate flap-type, single-stage linkage flap-type, and dual-stage linkage flap-type configurations. The separate flap-type configuration was tested through ground experiments. Control experiments were conducted under three different experimental conditions as follows: deflection of the upper cascades only, deflection of the lower cascades only, and symmetrical deflection of both cascades. For each condition, the deflection angles tested were 0°, 10°, 20°, 30°, 40°, 50°, and 60°. The thrust reductions observed for these three conditions were 0%, 37.5%, and 27.5% of the maximum thrust, respectively, without additional changes in the pitch moment. A combined thrust adjustment method maintaining a zero pitch moment demonstrated a linear thrust reduction to 20% of its initial value. The experiment concluded that the novel thrust control configuration effectively adjusted thrust without altering the BLI fans’ rotation speed, solving the coupled lift–thrust problem and enhancing BWB landing stability.

Analysis of blended winglet parameters on the aerodynamic characteristics of NXXX aircraft using Computational Fluid Dynamics (CFD)

Winglet is a wingtip device that was developed to improve aircraft flight performance by reducing induced drag. The induced drag produced by the NXXX aircraft is considered too excessive because the simple wingtip currently available is still not optimal in reducing induced drag, so a new more optimal wingtip is needed. This research aims to analyze the influence of blended winglet parameters on the aerodynamic characteristics of the NXXX aircraft such as CL and CD. Not only that, pitch moment coefficient (CM) is also studied to determine how trim drag can increase with the consequent increase of aerodynamic efficiency (CL/ CD) which currently has not been specifically researched yet. Winglet parameters in this study include variations in taper ratio, winglet height, cant angle, trailing edge sweep, and blending radius that there is not enough analysis about it now. Computational Fluid Dynamics (CFD) simulation uses Ansys CFX with the Shear Stress Transport (SST) turbulence model is used to obtain wing aerodynamic data. This research concludes that the best winglet configuration is taper ratio of 0.2, blending radius of 15 %, winglet height of 30 %, cant angle of 15°, and trailing edge sweep of −0.6°. Apart from that, it was also found that all winglet configurations increased CL, CLmax, and also lift slope (a). Almost all winglets also reduce the critical angle of attack, reduce CD, and increase CM. This simulation finds that addition of blended winglet can increase aerodynamic efficiency up to 17.51 % in cruise condition using the variation of winglet height.

Aerodynamic Characterization of a Hypersonic Projectile Using Acceleration-Based Measurements and Numerical Methods

Armor-piercing fin-stabilized projectiles stand out by their exceptionally high slenderness ratios and ground-level flight at super- and hypersonic speeds. As space constraints limit the integration of measurement equipment into such slender test models, nonintrusive measurement techniques become favorable. The present analysis demonstrates a renewed approach to the free-flight technique, which sets models into unconstrained flight through the test section. It enabled the evaluation of the quasi-steady drag, lift, and pitching moment characteristics, and, to some extent, also the dynamic pitch damping characteristics. An alternative to the free-flight technique is the free-oscillation technique, which limits the motion to a rotation around the center of gravity. The free-oscillation technique allowed a higher-fidelity analysis of the static and dynamic pitching moments. Both methods were based on the analysis of the accelerations derived from trajectory tracking. This acceleration-based approach enabled the evaluation of the highly nonlinear characteristics. The high slenderness of the investigated projectile leads to a significant contribution of the viscous forces to the overall drag. These viscous forces are sensitive to the laminar-turbulent boundary-layer state, as well as the thermal boundary conditions. The application of a high-resolution schlieren system enabled the assessment of the local laminar-turbulent boundary-layer state, while the use of low- and high-enthalpy testing facilities enabled the assessment of the aerothermal influence. Steady-state Reynolds-averaged Navier–Stokes simulations, which included laminar-turbulent transition modeling, were employed to replicate the results of the experimental efforts.

Experimental investigation on the hydrodynamic performance of a bioinspired manta-ray underwater vehicle in various forward propulsion modes

Bionic propulsion technology is revolutionizing underwater exploration by enabling underwater vehicles to navigate the oceans more efficiently. This research developed a scaled-down (10:1) experimental prototype of bionic manta-ray underwater vehicle, which has two degrees of freedom (2-DOF) pectoral fins: flapping and rotating. The impact of four different pectoral fin motion patterns (sinusoidal flapping and rotating, flapping frequency asymmetric, flapping amplitude offset, and rotating amplitude limitation) on the forward propulsion performance was investigated experimentally in a static water tunnel using a force/torque sensor and power analyzer. The results revealed that introducing asymmetry into the fin's flapping motion (asymmetric frequency or amplitude) significantly improved both average thrust and propulsive efficiency (the average thrust can be increased by more than 30%). Flapping with amplitude offset can effectively adjust its pitching moment without sacrificing much thrust. While the rotating motion of the pectoral fins contributes minimally to thrust generation compared to flapping, it plays an important role in enhancing both longitudinal stability and propulsive efficiency. Understanding the hydrodynamic performance of these various forward propulsion modes provides valuable insights for optimizing control strategies of bionic manta-ray underwater vehicles.

Numerical Analysis on The Aerodynamic Performance of A High-speed Train Operating on Various Embankment Heights Under The Influence of Crosswinds

Given the unavoidable geographical surface, railings must be raised above ground level in some cases, which is known as an embankment. It was discovered that the height of the embankment had a significant influence on the slipstream on the train's leeward side especially during crosswind conditions. The primary objectives of this study were to investigate the impact of varying embankment heights on the aerodynamic characteristics of a high-speed train under different crosswind conditions using computational fluid dynamics (CFD) analysis. The German Aerospace Center DLR's Next-Generation High-speed Train (NG-HST) model has been used for this study. The yaw angles (ψ) are ranging from 10° to 50° in 10° increments. The Reynolds number based on the model's height and freestream velocity at the computational domain is 1.3 x 106. In the results, it shows that embankment height and crosswind ψ has a significant impact on the aerodynamic characteristics. Essential aerodynamic parameters that have a significant impact on train stability, such as the drag, lift, and side force coefficients, as well as the roll, yaw, and pitch moment coefficients, revealed that the higher the ψ, which included 40° and 50°, produced poor results compared to the lower ψ, which included 10°, 20°, and 30°. In terms of visual appearance, rising of crosswind angles have a greater impact on the formation of vortices on the leeward side of the train body and embankment. Thus, it can be concluded that the embankment heights and crosswind angles are crucial in determining train safety operations.

Numerical simulation of the transition flight aerodynamics of cross-shaped quad-tiltrotor UAV

In order to enhance the stability of the tilt transition process, a new configuration of Quad-Tiltrotor UAV was presented in this paper. Firstly, numerical simulation was used to calculate and analyze the aerodynamic interaction between the front rotor/fuselage/rear rotor during the transition state mode. The calculation model of the isolated rotor, front-rear rotor, front rotor-fuselage, and front rotor-rear rotor-fuselage combination states are established. Besides, the effects of pitch, roll, and yaw moment on the fuselage at different tilt angles are analyzed. It is concluded that the front rotor is the leading factor in the aerodynamic interference of the whole UAV in the different combination states. The research results can provide a reference for the optimization design of the overall layout, structure, and flight control strategy of the cross-shaped quad-tiltrotor UAV, and can also provide solutions for the logistics application of urban air traffic.

Efficient computation of horizontal stabilizer incidence angle for low pitching moments based on actuator model

Efficient computation of horizontal stabilizer incidence angle for low pitching moments based on actuator model

Control of Helicopter Using Virtual Swashplate

This article presents a virtual swashplate mechanism for a mini helicopter in classic configuration. The propeller bases are part of a passive mechanism driven by main rotor torque modulaton, this mechanism generates a synchronous and opposite change in the propellers angle of attack, then the thrust vector tilts. This approach proposes to control the 6 degrees of freedom of the aircraft using two rotors. The main rotor controls vertical displacement and uses torque modulation and swing-hinged propellers to generate pitch and roll moments and the horizontal displacement while the yaw moment is controlled by the tail rotor. The dynamic model is obtained using the Newton-Euler approach and robust control algorithms are proposed. Experimental results are presented to show the performance of the proposed virtual swashplate in real-time outdoor hover flights.

Aerodynamic Feedforward-Feedback Architecture for Tailsitter Control in Hybrid Flight Regimes

This paper presents a guidance-control design methodology for the autonomous maneuvering of tailsitter transitioning unmanned aerial systems (t-UASs) in hybrid flight regimes (i.e., the dynamics between VTOL and fixed-wing regimes). The tailsitter guidance-control architecture consists of a trajectory planner, an outer-loop position controller, an inner-loop attitude controller, and a control allocator. The trajectory planner uses a simplified tailsitter model, with aerodynamic and wake effect considerations, to generate a set of transition trajectories with associated aerodynamic force estimates based on an optimization metric specified by a human operator (minimum time transition). The outer-loop controller then uses the aerodynamic force estimate computed by the trajectory planner as a feedforward signal alongside feedback linearization of the outer-loop dynamics for 6DOF position control. The inner-loop attitude controller is a standard nonlinear dynamic inversion control law that generates the desired pitch, roll, and yaw moments, which are then converted to the appropriate rotor speeds by the control allocator. Analytical conditions for robust stability are derived for the outer-loop position controller to guarantee performance in the presence of uncertainty in the feedforward aerodynamic force compensation. Finally, both tracking performance and stability of the control architecture are evaluated on a high-fidelity flight dynamics simulation of a quadrotor biplane tailsitter for flight missions that demand high maneuverability in transition between flight modes.

Effects of the gap ratio on the flow field structures and the aerodynamic performance of an airfoil with ridge ice

Under SLD icing conditions, the ridge ice may appear on the surface of aircraft, which led to the significant aerodynamic deterioration and affected aircraft flight safety. The present study experimentally investigated the effects of the gap ratio on the flow field structures and aerodynamic performance of an airfoil with ridge ice. Detailed measurements were performed in a low-speed reflux wind tunnel by utilizing the Particle Image Velocimetry (PIV) technique and a high-sensitivity six-component balance. The results showed that the maximum lift coefficient, stall angle, and maximum pitch moment coefficient of the airfoil increased as the gap ratio enlarged. At AOA = 10 deg, the separation bubble length decreased by 77 % when the gap ratio changed from 0 to 0.1. Meanwhile, the separation bubble length decreased by 68 % when the gap ratio changed from 0.1 to 0.3. Besides, as the increase of the gap ratio, the average vorticity, turbulent kinetic energy, and Reynolds shear stress in the selected region above the airfoil decreased, while the average velocity increased. In addition, the gap ratio did not have an apparent effect on the transition onset positions and the maximum spanwise vorticity in the flow field.

Comparative study of nonlinear wave-induced global loads on a container ship in regular head waves predicted with numerical methods based on weakly nonlinear potential theory and Reynolds-averaged Navier-Stokes equations

ABSTRACT Wave-induced loads on the ship hull are of significant interest for many reasons. These loads, however, are highly nonlinear in realistic wave conditions and consequently difficult to predict. In this paper, we analysed and compared wave-induced loads acting on a container vessel in regular head waves computed with a newly developed weakly nonlinear potential theory-based code and a flow solver based on the Reynolds-averaged Navier-Stokes Equations. We investigated both long and short waves with moderate steepness. Initially, the vessel was completely fixed and the ship motions were suppressed, then the vessel was free to heave and pitch motion. We compared time-histories of the computed wave-induced longitudinal and vertical global forces as well as the pitch moment and we deeply analysed nonlinear effects by applying a Fourier transformation and comparing the corresponding harmonic amplitude up to the fourth order. Our results demonstrated satisfactory agreement between the two numerical methods, especially when the ship was free to heave and pitch. Additionally, the results provided detailed information about the capabilities and limitations of the weakly nonlinear potential theory-based approach.

High-Order Wall-Modeled Large-Eddy Simulation of High-Lift Configuration

This paper presents the assessment of several recent enhancements for a high-order wall-modeled large-eddy simulation (WMLES) approach and demonstrates order independence with a fixed data exchange location in the wall model. The two enhancements include the use of isotropic tetrahedral elements to improve accuracy and an explicit subgrid-scale model, the Vreman model, to improve accuracy and robustness. The P-refinement study focused on the high-lift Common Research Model (HL-CRM) at the angle of attack of 19.57 deg, a benchmark problem from the 4th AIAA High-Lift Prediction Workshop. Solution polynomial orders of P=2, 3, 4, and 5 were used in the study. The study demonstrated p independence in integrated forces, pitch moment, velocity profile in the wall-normal direction, and surface flow topology. It also showed that a P order of at least 3 (P3) was needed to correctly predict the external inviscid flow and the surface flow topology. Thereafter, P3 simulations over several other angles of attack demonstrated that the high-order WMLES approach can correctly predict the maximum lift and flow separation regions for HL-CRM with about 40 million degrees of freedom (DOF) compared to at least 250 million DOF required by second-order methods.

Static and dynamic characteristics of supersonic cruise missile with damaged wing

AbstractAccurately evaluating the aerodynamic performance of the missile with damaged structures is very important for the subsequent flight control strategy. At present, few researchers have studied the aerodynamic characteristics of damaged supersonic cruise missiles. Based on CFD (computational fluid dynamics) solutions and the dynamic derivative identification method, the differences in static and dynamic characteristics between the damaged and undamaged models are compared. The results indicate that when the extent of damage increases, the change rate of drag coefficient at larger AoA (angle-of-attack) is greater than that at the smaller AoA. On the contrary, the change rate of lift coefficient at larger AoA is smaller than that at smaller AoA. Meanwhile, the absolute value of the static pitch moment decreases, but the absolute value of the roll moment increases. Damage causes a change in the absolute values of the pitch and roll dynamic derivatives, and the dynamic derivatives do not vary monotonically with the increase of AoA. The turning point occurs at about $\alpha$ = 5°. The areas of the hysteresis loops of the pitch-roll coupling moment increase, which makes the dynamic coupling characteristic between the pitch and roll directions increase. Finally, the maximum allowable damage extent of the missile wing that can achieve static trim is obtained and validated by controlling the deflection of the four rudders.

Experimental and Numerical Investigations of Mini-Tab Device for Active Flow Control

The next generation of aircraft is likely to exhibit a breadth of aeromechanic and aeroelastic behaviors that will be exacerbated by the turbulent flows generated in urban environments. Exploitation of novel flow control technologies will be critical in offering greater control authority to enable this new generation of aerial vehicles for faster, safer, and more efficient flight. This study experimentally investigates the efficacy of a mini-tab active flow control device using a novel dynamic test rig. Under steady conditions, the mini-tab response exhibited nonlinearity with deployment height. Under dynamic conditions, the lift and pitching moments displayed hysteresis around their quasi-steady counterparts at low reduced frequencies (k=0.05) but displayed significant deviations at middle to high reduced frequencies (k=0.15 to 0.31) due to the formation of a tab vortex that propagates downstream. It was demonstrated, for the first time experimentally, that mini-tab dynamic performance is insensitive to unsteady airfoil motions in pitch and plunge—representative of flutter onset. Finally, an existing mini-tab model was extended to capture vortex formation on the upper and lower airfoil surfaces and modified to incorporate mini-tab rate-dependence on vortex initiation. The model showed good agreement with experimental data in both periodic and transient scenarios.

Integrated Three-Dimensional Airloads and Stresses on Lift-Offset Coaxial Rotors at Extreme Speeds

This paper investigates the three-dimensional (3-D) dynamic stresses on a modern four-bladed hingeless coaxial rotor—inspired by the gross dimensions of the Sikorsky S-97 Raider at extreme flight speeds. The stresses are obtained using integrated 3-D (I3D) aeromechanical analysis—defined as the coupling of 3-D finite element-based structural dynamics with 3-D Reynolds-averaged Navier–Stokes–based fluid dynamics. The coupling was carried out with the University of Maryland/U.S. Army structural dynamic solver X3D and the U.S. Army CREATETM–AV Helios suite of fluid dynamic solvers. The new structural analysis is both enabled and driven by advanced high-performance computing, parallel and scalable solvers, high-order 3-D brick finite elements unified with multibody dynamics, integrated aeromechanics, and a special 3D-to-1D fluid–structure interface that refines the power of the delta-coupling procedure while retaining the advantages of existing computational fluid dynamics mesh motion schemes. The analysis is carried out at 220 knots (μ=0.5)—the cruise speed of the S-97 Raider without reduced tip speed—in order to study the stresses in extreme conditions. At such high speeds, the blade lift is dominated by the complex tip vortex roll-up, and the pitching moments and drag are dominated by the unsteady transonic shocks at the tip. Interesting 3-D dynamic stress patterns are revealed all across the blade that have remained invisible until now since they could neither be predicted nor measured in flight. The key conclusion is that such high-fidelity analysis is now indeed possible and, in fact, necessary to get deeper insights into the dynamics of coaxial rotors at extreme speeds.

Liquid sloshing in a tank vehicle: A numerical and experimental investigation with baffle configurations

A three-dimensional numerical model based on VOF (Volume of Fluid) method is established to investigate anti-sloshing effects of various baffle configurations, proposed in this paper, on liquid sloshing in a partially filled tank vehicle. A series of experiments were conducted in a scaled clean-bore or baffled tank. The liquid surface deformations obtained from the numerical model exhibit good agreement with the experimental results at each moment, demonstrating the validity of the numerical model. The validated numerical model was further used to investigate the anti-sloshing effect of different baffle configurations in full-size tank vehicles. The results imply that the addition of baffles can decrease longitudinal force and pitch moment to some extent, in terms of peak value or mean value. Moreover, the addition of baffles may elevate the fundamental sloshing frequency by 50% to 200%, mitigating the risk of resonance phenomena in tank vehicles.

Improving the Hydrodynamic Performance of Underwater Tags for Blue Shark Monitoring

The use of tag devices in marine environments has become indispensable in attaining a better understanding of marine life and contributing to conservation efforts. However, the successful deployment and operation of underwater tags both depend significantly on their hydrodynamic characteristics, particularly their resistance to motion and stability in various environmental conditions. Herein, a comprehensive study on the hydrodynamic characteristics and optimization of an underwater tag designed for monitoring blue sharks is presented. Firstly, a validation process is conducted by comparing the computational fluid dynamics (CFD) results with the experimental data from Myring’s study, focusing on the resistance characteristics of the tag’s body and the impact of various operational conditions. Subsequently, the validated CFD model is applied to assess the hydrodynamic performance of the tag under different flow conditions, velocities, and angles of attack. Through iterative simulations, including mesh independence studies and boundary condition adjustments, the study identifies key parameters influencing the tag’s resistance and stability. Furthermore, the paper proposes and implements design modifications, including the incorporation of stabilizing fins, aimed at minimizing resistance and improving the tag’s equilibrium position. The effectiveness of these design enhancements is demonstrated through a comparative analysis of resistance and pitching moments for both preliminary and optimized tag configurations. Overall, the study provides valuable insights into the hydrodynamic behavior of underwater tags and offers practical recommendations for optimizing their design to minimize interference with the movement of tagged marine animals.

Indirect load measurement method and experimental verification of floating offshore wind turbine

Accurate load data of wind turbines operating in real sea conditions is crucial for real-time control and offline structural optimization. However, the sheer size of these structures and the complexities of their working environments challenge the direct load measurement through strain gauges and fiber Bragg grating sensors. To address these problems, this paper introduces an indirect measurement method to acquire the thrust and pitch moment loads of the turbine. First, load identification techniques based on the frequency domain separation and aerodynamic damping were applied to decouple the wind and wave loads and establish the inverse load calculation model. Then, numerical simulations were carried out to calibrate and verify this model under a range of steady and unsteady conditions. Preliminary conclusions indicated the strong viability of this method, revealing a mean relative error of 2.32% for thrust and a corresponding equivalent high-frequency error of 24.66% for pitch moment under the designed operating conditions. Finally, a hybrid scaled model was constructed, combining a virtual wind turbine with a physical platform. Experimental results in the wave basin demonstrated the effectiveness of this indirect method in measuring thrust loads of wind turbines, achieving a mean relative error of less than 10%.

Dynamic characteristics and application of dual throat fluidic thrust vectoring nozzle

Thrust vectoring technology plays an important role in improving the maneuverability of aircraft. In order to overcome the disadvantages of mechanical thrust vectoring nozzles, such as complications of structure and significant increases in weight and cost, fluidic thrust vectoring nozzles are proposed. Dual Throat fluidic thrust vectoring Nozzle (DTN) has received wide attention due to its excellent thrust vectoring efficiency and minimal thrust loss. In this study, three-dimensional unsteady numerical simulations of a single axisymmetric DTN are conducted first to analyze its dynamic response. Then the pitch and yaw control characteristics of DTN equipped on a flying-wing aircraft are investigated. It is found that the dynamic response will experience three stages: rapid-deflecting stage, oscillating stage, and steady stage. A complete recirculation zone forms at the end of the rapid-deflecting stage, which pushes the primary flow to attach to the wall opposite the secondary injection. Meanwhile, the exhaust flow is deflected. In terms of DTN’s application, the DTN equipped on the flying-wing aircraft is capable of providing effective pitch and yaw moments at all angles of attack and Mach numbers. In addition, continuous pitch and yaw moments can be obtained by adjusting the secondary mass flow ratios. The control moment is generated due to the asymmetrical pressure distribution of nozzle surface, which is mainly contributed by the pressure decrease on the secondary injection surface. Moreover, the DTN equipped on the flying-wing aircraft has a relatively high thrust vectoring efficiency of around 5°/% and a thrust coefficient of around 0.95 when nozzle pressure ratio equals 4. These results provide an important theoretical basis for the practical application of DTN.