Academic literature on the topic 'Airspeed reduction'

Vultures are thought to form networks in the sky, with individuals monitoring the movements of others to gain up-to-date information on resource availability. While it is recognized that social information facilitates the search for carrion, how this facilitates the search for updrafts, another critical resource, remains unknown. In theory, birds could use information on updraft availability to modulate their flight speed, increasing their airspeed when informed on updraft location. In addition, the stylized circling behaviour associated with thermal soaring is likely to provide social cues on updraft availability for any bird operating in the surrounding area. We equipped five Gyps vultures with GPS and airspeed loggers to quantify the movements of birds flying in the same airspace. Birds that were socially informed on updraft availability immediately adopted higher airspeeds on entering the inter-thermal glide; a strategy that would be risky if birds were relying on personal information alone. This was embedded within a broader pattern of a reduction in airspeed (approx. 3 m s −1 ) through the glide, likely reflecting the need for low speed to sense and turn into the next thermal. Overall, this demonstrates (i) the complexity of factors affecting speed selection over fine temporal scales and (ii) that Gyps vultures respond to social information on the occurrence of energy in the aerial environment, which may reduce uncertainty in their movement decisions.

This paper examines rotor power reductions achievable through a combination of radius and RPM variation. The study is based on a utility helicopter similar to the UH-60A and considers +17% to –16% variation in radius and ±11% variation in RPM about the baseline, over a range of airspeed, gross weight, and altitude. Results show that decreasing RPM alone effectively reduced power at cruise velocities in low-and-light conditions, but the power reductions diminished at increasing altitude and/or gross weight, and in low-speed flight. Increasing radius alone, on the other hand, had greatest effectiveness in power reduction in high-and-heavy operating conditions and at lower flight speeds. When radius and RPM variation is used in combination, minimum RPM is always favored, along with radius increases at increasing altitude and gross weight, and in low-speed operation. At low-to-moderate gross weight, the significant power reductions seen in cruise and at low altitude with RPM variation alone are obtained even at higher altitude, and over the airspeed range, using radius and RPM variation in combination. In high-and-heavy conditions, the combination of RPM reduction and radius increase yields very large power reductions of over 20% and up to 30% over the baseline. Power reduction in low-and-light conditions comes almost entirely from profile power reduction due to RPM decrease. In cruise and high-speed flight, the profile power reductions progressively give way to induced power reductions at increasing gross weight and altitude. At low speeds, reduction in induced power due to increased radius and decreased disk loading dominates.

Abstract. Drift is one of the most hazardous consequences of an improper aerial application of glyphosate. Wind, droplet size, application height, and distance to sensitive areas are the most important factors for drift. Droplet size is affected by nozzle, operating pressure, flight speed, deflection angle, and physicochemical properties of the spray solution. The objective of this study was to evaluate the effect of flight speed and the use of adjuvants on droplet size spectra in aerial applications of glyphosate. The study was conducted in a high-speed wind tunnel at the Pesticide Application Technology Laboratory (University of Nebraska-Lincoln, West Central Research and Extension Center, North Platte, Neb.). Aerial applications were simulated with four different airspeeds (44.4, 52.8, 61.1, and 69.4 m/s) and glyphosate combined with adjuvants (high surfactant oil concentrate, microemulsion drift reduction agent, nonionic and acidifier surfactant, polyvinyl polymer, and glyphosate alone). Droplet size spectra were evaluated using a Sympatec Helos laser diffraction instrument measuring 90 cm from the nozzle tip (CP11-4015). The volumetric droplet size distribution parameters (VMD, DV0.1, and DV0.9) and the percentage of droplets smaller than 100 µm were reported. The relative span was calculated to indicate the droplet size homogeneity [(DV0.9 - DV0.1) / DV0.5]. Glyphosate solutions with adjuvants had a larger VMD than the glyphosate alone solution at 44.4 m/s wind speed. At 69.4 m/s only the glyphosate solution with polymer had a larger VMD. Conversely, the glyphosate with polymer had the smallest DV0.1, and the greatest relative span and percentage of droplets smaller than 100 µm. Generally, adjuvants influence on droplet size was diminished or muted as the airspeed was increased. The polymer tested in this study failed as a drift agent reduction agent, especially at higher airspeeds. While not all polymers were tested, cautions should be taken if using these types of adjuvants in aerial applications. The interaction of airspeed and adjuvants influencing droplet size distribution in aerial applications of glyphosate should be considered by applicators in order to mitigate glyphosate drift to the surrounding environment. Further studies are necessary to better understand the interaction between solution viscosity and air shear effect on the atomization process and droplet size distribution, as well as confirm that trends hold true for other adjuvants in the polymer class. Although applicators tend to operate aircrafts with increased flight speeds in order to optimize the application time efficiency, this practice can reduce or mute adjuvants effects, decrease the droplet size distribution, and increase drift potential in aerial applications of glyphosate. Keywords: Drift reduction technologies, Flight speed, High-speed wind tunnel, Laser diffraction.

The rapid growth of air travel and aviation emissions in recent years has contributed to an increase in climate impact. Contrails have been considered one of the main factors of the aviation-induced climate impact. This paper deals with the formation of persistent contrails and its relationship with fuel consumption and flight time when flight altitude and true airspeed vary. Detailed contrail formation conditions pertaining to altitude, relative humidity and temperature are formulated according to the Schmidt–Appleman criterion. Building on the contrail formation model, the proposed model would minimise total travel time, fuel consumption and contrail length associated with a given flight. Empirical data (including pressure, temperature, relative humidity, etc.) collected from seven flight information regions in Chinese observation stations were used to analyse the spatial and temporal distributions of the persistent contrail formation area. The trade-off between flight time, fuel consumption and contrail length are illustrated with a real-world case. The results provided a valuable benchmark for flight route planning with environmental, flight time, sustainable flight trajectory planning and fuel consumption considerations, and showed significant contrail length reduction through an optimal selection of altitude and true airspeed.

An aeroservoelastic system can be described as a gridding-based linear parameter-varying (LPV) model, whose dynamic characteristics usually vary with the airspeed. Due to the high order of the system, it is necessary to perform order reduction on LPV models to overcome the control design challenges. However, when directly extending linear time-invariant (LTI) model order reduction technologies to the LPV system, states of the reduced-order LTI models generated separately at different grid points could be inconsistent. In this paper, a novel modal matching algorithm is proposed to solve the problem of state inconsistency by identifying the internal connection between the models at adjacent grid points. An oblique projection-based distance metric is defined to improve the reliability of the modal matching algorithm. The reduced-order LPV model constructed based on this method would have a high fidelity relative to the original model and a smooth interpolation performance between grid points. The proposed algorithm is applied to the X-56A aircraft, and numerical results show its effectiveness.

In this paper, a small airplane is redesigned by using a distributed electrical propulsion (DEP) system. The design procedure is focused on the reduction of fuel consumption in cruise regime with constrained parameters of take-off/landing. In this case, a one half wing area compared to an original airplane is used. Take-off distance and minimum airspeed for landing is achieved by distributed propellers mounted on the leading edge of the wing. These propellers induce velocity on the wing and thereby increase local dynamic pressure, thus the required lift force can be reached with smaller wing area. Moreover, the distributed propellers are assumed as folded in cruise regime to minimize drag when the main combustion engine provides sufficient power.

ABSTRACTAir traffic flow becomes denser and more complex within terminal manoeuvering areas (TMAs) due to rapid growth rates in demand. Effective TMA arrival management plays a key role in the improvement of airspace capacity, flight efficiency and air traffic controller performance. This study proposes a mixed integer linear programming model for aircraft landing problems with area navigation (RNAV) route structure using three conflict resolution and sequencing techniques together: flexible route allocation, airspeed reduction and vector manoeuver. A two-step mixed integer linear programming model was developed that minimises total conflict resolution time and then total airborne delay using lexicographic goal programming. Experimental results demonstrate that the model can obtain conflict-free and time optimal aircraft trajectories for RNAV route structures.

The high cost of conventional energy production, its uses, and its negative environmental impact led scientists to look for many ways to solve this problem. The main issues are environmental pollution, the high cost of air treatment, and the high cost of energy production. An essential solution is using renewable energy instead of traditional energy (petroleum, coal, natural gas, and electric power). As it is a permanent energy source, it causes no air pollution and is cheaper. In this work, an experimental study was conducted on using solar energy as an energy source in the natural ventilation of buildings and the factors affecting this ventilation. A solar chimney has been installed with and without storage materials to store energy for natural identity during sunrise and after sunset. In this work, the difference in air density with different temperatures is the main factor for energy production, which leads to a change in the airspeed (natural ventilation). The study indicates that indoor air velocity can remain higher than ambient air velocity during the day and at night for several hours. This reduces demand and dependence on conventional energy usage, which leads to cost reductions and a reduction in air and environmental pollution.

Turbulent atmospheric conditions represent a challenge to stable flight in soaring birds, which are often seen to drop their wings in a transient motion that we call a tuck. Here, we investigate the mechanics, occurrence and causation of wing tucking in a captive steppe eagle Aquila nipalensis , using ground-based video and onboard inertial instrumentation. Statistical analysis of 2594 tucks, identified automatically from 45 flights, reveals that wing tucks occur more frequently under conditions of higher atmospheric turbulence. Furthermore, wing tucks are usually preceded by transient increases in airspeed, load factor and pitch rate, consistent with the bird encountering a headwind gust. The tuck itself immediately follows a rapid drop in angle of attack, caused by a downdraft or nose-down pitch motion, which produces a rapid drop in load factor. Positive aerodynamic loading acts to elevate the wings, and the resulting aerodynamic moment must therefore be balanced in soaring by an opposing musculoskeletal moment. Wing tucking presumably occurs when the reduction in the aerodynamic moment caused by a drop in load factor is not met by an equivalent reduction in the applied musculoskeletal moment. We conclude that wing tucks represent a gust response precipitated by a transient drop in aerodynamic loading.

An efficient sintering process was proposed based on the autocatalytic denitrification of the sintered ore. The catalytic denitrification of sintered ore, the effect of double-layer ignition sintering process on the emission reduction in nitrogen oxides, and the impact on the quality of sintered ore were studied. The results showed that the catalyzed reduction of NO with sinter ore as a catalyst has a significant effect; when the airspeed reaches 3000 h−1, the temperature is 500 °C, and the conversion rate of NO can reach 99.58%. The sinter yield of double-layer ignition sintering is increased, solid fuel consumption is slightly reduced, falling strength is slightly increased, and drum strength is slightly decreased. Under the conditions of layer height proportion of 320/400 mm (lower/upper) and ignition time interval of 10 min, the yield, drum strength, shatter strength, and solid fuel consumption reached 61.60%, 54.82%, 46.75%, and 69.55%, respectively. NOx concentration under the 16% baseline oxygen content (c(NOx)’) in the flue gas of double-layer ignition sintering is reduced to a certain extent, and the generation time of NOx is greatly shortened. The double-layer ignition sintering process can reduce the emission of nitrogen oxides in the sintering process under the condition of guaranteeing the quality of sinter, which has great economic and environmental benefits.

Le projet de thèse consiste à développer une solution pour l'atterrissage d'un drone à voilure fixe de configuration classique dans une zone limitée. Le principal défi consiste à réduire la vitesse de l'avion à une phase minimale pendant le vol, à l'aide d'algorithmes de contrôle automatique. La réduction de la vitesse d'un drone à voilure fixe s'effectue en augmentant son angle d'attaque, ce qui implique un freinage par la force de traînée. Cependant, cette manœuvre est critique pour un avion conventionnel, parce que si son angle d'attaque augmente au-delà de l'angle de décrochage, le véhicule peut perdre sa contrôlabilité, c'est-à-dire qu'il est possible que le véhicule aérien s'effondre et que sa structure soit endommagée. Le modèle mathématique est une représentation d'équations qui décrit le comportement de la dynamique du système. En considérant plusieurs variables pour obtenir une meilleure approximation de la dynamique du système, dans notre cas le véhicule à voilure fixe, la conception des stratégies de contrôle sera plus difficile et plus complexe. Dans ce travail de recherche, nous utiliserons un modèle mathématique non linéaire car les effets de décrochage peuvent être inclus par des approximations mathématiques du moment de tangage, des forces de portance et de traînée. Cela nous permet d'obtenir une meilleure performance des lois de contrôle pour la navigation autonome du drone à voilure fixe. L'une des limites des véhicules à voilure fixe est qu'ils atterrissent dans des espaces de dimensions réduites et que le pourcentage de dommages subis par leur structure est élevé. En outre, les perturbations extérieures et l'inexpérience des pilotes augmentent le risque de dommages. Il est bien connu qu'il est très difficile de satisfaire aux conditions d'une piste d'atterrissage. Par conséquent, la communauté scientifique s'est efforcée de mettre au point des solutions pour l'atterrissage dans des zones limitées. Dans la littérature, on trouve quelques solutions basées sur des véhicules hybrides et des systèmes de récupération. Les véhicules hybrides consistent à modifier la structure d'un véhicule à voilure fixe. Les moteurs sont répartis stratégiquement pour obtenir une configuration de véhicule multirotor, offrant certaines caractéristiques telles que le décollage et l'atterrissage verticaux. Cependant, ces actionneurs augmentent la masse du véhicule, la consommation d'énergie (ce qui réduit la durabilité du vol), la probabilité de défaillance, le coût d'acquisition, de réparation et d'entretien. Notre objectif dans ce travail de recherche est de concevoir et de valider des stratégies de contrôle pour l'atterrissage d'un drone à voilure fixe dans un espace limité. Les stratégies de contrôle ont été conçues selon deux approches : la première est basée sur le développement de manœuvres pour un drone à voilure fixe afin de réduire la vitesse à une phase minimale pendant le vol. Dans la deuxième approche, nous avons travaillé sur les stratégies de contrôle pour l'atterrissage d'un drone à voilure fixe sur un véhicule terrestre en mouvement. Une stratégie de contrôle a été proposée pour réduire la vitesse du drone à voilure fixe au minimum afin d'être capturé par un système de récupération. La stratégie de contrôle a été divisée en trois étapes de vol : dans la première étape, l'avion s'aligne dans le plan x-y tandis qu'il est conduit à une altitude souhaitée pour effectuer un vol de croisière. L'étape suivante consiste en un vol ascendant, axé sur le suivi d'une référence angulaire basée sur une trajectoire phugoïde. Cette trajectoire implique une augmentation de l'angle d'attaque jusqu'à l'angle de décrochage de l'avion. Ainsi, la vitesse aérienne obtient une réduction maximale dans des conditions sûres, permettant au drone d'être capturé par le système de récupération. Toutefois, si le drone n'est pas capturé par le système de récupération, une stratégie de contrôle est appliquée pour rétablir le vol de l'aéronef
The development of this thesis consists of designing some control strategies that allow a fixedwing drone with classical configuration to perform a safe landing in a limited area. The main challenge is to reduce the aircraft’s airspeed avoiding stall conditions. The developed control strategies are focused on two approaches: the first approach consists of the designing airspeed reduction maneuvers for a fixed-wing vehicle to be captured by a recovery system and for a safe landing at a desired coordinate. The next approach is focused on landing a fixed-wing drone on a moving ground vehicle. A dynamic landing trajectory was designed to lead a fixedwing vehicle to the position of a ground vehicle, reaching its position in a defined distance. Moreover, this trajectory was used in a cooperative control design. The control strategy consists of the synchronization of both vehicles to reach the same position at a desired distance. The aerial vehicle tracks the dynamic landing trajectory, and the ground vehicle controls its speed. In addition, we will propose a control architecture with a different focus, where the ground vehicle performs the tracking task of the aerial vehicle’s position in order to be captured. And, the drone’s task is to track a descending flight until the top of the ground vehicle. However, considering the speed difference between both vehicles. Therefore, we propose a new control architecture defining that the aircraft performs an airspeed reduction strategy before beginning its landing stage. The aircraft will navigate to a minimum airspeed, thus, allowing the ground vehicle to reach the fixed-wing drone’s position by increasing its speed. The control laws of each strategy were determined by developing the Lyapunov stability analysis, thus, the stability is guaranteed in each flight stage. Finally, the control strategies were implemented on prototypes allowing us to validate their performance and obtain satisfactory results for safe landing of a fixed-wing drone with classical configuration

FAA rotorcraft airworthiness regulations require calibration of pitot-static systems in all flight regimes. Of all methods commonly used, none has been applied in a manner showing full compliance, specifically in the takeoff phase and in determining CG (Center of Gravity) effects. A review of accepted Position Error Correction methods identifies the GPS-based true airspeed method, with an adapted execution and analysis technique, as the most practical in terms of equipment and efficiency to provide a complete airspeed system calibration. The level flight limitations of the GPS method are solved by a combination of flight profiles, continuous data recording and reduction technique. The GPS horseshoe method and the ORBIS constant turn radius method are expanded by varying the airspeed, altitude, and heading as required to provide an equation set solved for the wind components and true airspeed. The new variable parameter methods minimize wind variability effects and flight test time.

Autorotation is an emergency maneuver executed by helicopter pilots usually following a loss of power, mechanical or system failure. The pilot is required to perform several tasks simultaneously and the timing of each of these must be precisely controlled. Workload can be high throughout the maneuver and the consequences of getting things wrong can be serious, if not fatal. Following on from a series of studies that investigated the use of both head-up displays and haptic cueing methods to assist pilots to fly the maneuver more safely and consistently, this paper presents a pilot-in-the-loop flight simulation study to explore the use of pilot cues provided on a head-down display. These provided pitch angle (and hence airspeed) cueing to the pilot. The cueing symbology was tested for a straight-in autorotation approach for a number of different visual environments and the results compared to equivalent cases where no cueing was present. The cues were assessed both subjectively, via pilots providing Bedford Pilot Workload Ratings, as well as objectively, via analysis of the flight path performance achieved. The subjective evaluation showed that cues using the head-down display were useful in terms of pilot workload reduction, particularly for the degraded visual environment scenario tested. The objective assessment revealed that pilots were able to establish the steady state descent more quickly and maintain forward airspeed within the specified bounds more reliably with the aid of the cues.

Emerging vertical flight concepts being proffered for solutions to the Future Vertical Lift (FVL) mission set such as compound high speed rotorcraft can be designed with multiple, coupled control effectors thus creating redundant systems in one or two more axes to generate control forces and moments which allow for a range of trim states. In the FVL mission area future rotorcraft will be asked to fly into high threat environments where potential failure modes can be encountered due to enemy fire or mechanical failure causing reduction of the safe flight envelope. Fault detection creates options to increase the survivability of the crew and passengers allowing an emergency flight envelope to be proposed. One of the more serious potential failures due to enemy fire is a loss of yaw control. Faults in yaw control can be detected in a compound rotorcraft with a vectored thrust ducted propeller (VTDP) or similar anti-torque thruster. An online Kalman filter (KF) for a dimensional yaw moment coeff icient model will be used to estimate vehicle yaw coeff icients. Deviation from the nominal coefficients will be monitored based on the KF statistics in the case of both rudder and tail rotor failure at 60, 40, and 20 ft/s in forward flight. Both frozen zero rudder and ganged sector faults as well as failed tail rotor faults were successfully detected at all airspeeds except the failed tail rotor at 60 ft/s. For the yaw control faults considered, post fault excitation appears airspeed dependent. An online KF estimator for yaw control fault detection could successfully be integrated into the design of a compound rotorcraft with VTDP thereby increasing system safety.

<div class="section abstract"><div class="htmlview paragraph">In this research, the performance of two commercially available icephobic coatings is evaluated on an 81% scaled-down version of the Bell Flight APT 70 drone propeller. Tests are performed in an icing wind tunnel (IWT) under selected severe icing conditions to test the ice protection capability of coatings against both glaze and rime ice. Two different coating formulations are used, one is a polydimethylsiloxane (PDMS) acetoxy terminated coating, the other an epoxy-silicone. The coatings were briefly characterized in terms of their surface roughness, water contact angle and ice adhesion reduction factor compared to aluminum using the centrifugal adhesion test (CAT). Blade sets were prepared for both coatings and a third uncoated set was tested for reference purposes. Tests in the IWT were performed to simulate a true airspeed of 35 m/s and a constant propeller rotational speed of 5 500 RPM. Two conditions of liquid water content (LWC) and droplet median volumetric diameter (MVD) were considered: LWC = 0.8 g/m³, MVD = 20 μm and LWC = 0.2 g/m³, MVD = 40 μm. The first condition was performed at static air temperatures of -5°C, -12°C and -20°C while the second was only performed at -5°C. The performance of the propeller is evaluated by means of the relative change in thrust coefficient, the torque coefficient and propeller efficiency. Tests were conducted such that operating conditions are maintained until vibration limits exceed the tolerated threshold to allow the possibility for multiple ice shedding events. Results demonstrated that the PDMS coating successfully reduced ice adhesion for all tests conditions while the epoxy-silicone only reduced ice adhesion for tests conditions at -5°C. The ice protection provided by either coating is shown to be insufficient to ensure safe flight under icing conditions due to significant propeller performance degradation and severe vibrations caused by non-symmetrical ice shedding.</div></div>

<div class="section abstract"><div class="htmlview paragraph">In-flight atmospheric icing is a severe hazard for propeller-driven unmanned aerial vehicles (UAVs) that can lead to issues ranging from reduced flight performance to unacceptable loss of lift and control. To address this challenge, a physics-based first principles model of an electric UAV propulsion system is developed and identified in varying icing conditions. Specifically, a brushless direct current motor (BLDC) based propeller system, typical for UAVs with a wing span of 1-3 meters, is tested in an icing wind tunnel with three accreted ice shapes of increasing size. The results are analyzed to identify the dynamics of the electrical, mechanical, and aerodynamic subsystems of the propulsion system. Moreover, the parameters of the identified models are presented, making it possible to analyze their sensitivity to ice accretion on the propeller blades. The experiment data analysis shows that the propeller power efficiency is highly sensitive to icing, with a 40% reduction in thrust and a 16% increase in torque observed on average across the tested motor speeds and airspeeds. The resulting reduction in propeller efficiency can be as high as 70% in the worst-case scenario. These findings provide valuable insights into the impact of ice accretion on electric propeller systems and contribute to the development of effective ice protection systems for safer UAV operation in cold environments.</div></div>

This paper addresses the use of dynamic inversion with direct load feedback to provide combined load alleviation and flight control of rotorocraft. The method is applied to a compound utility rotorcraft with similar airframe properties as a UH-60A along with a lifting wing. The controller makes use of flaperons and horizontal stabilizer in addition to the conventional main rotor / tail rotor blade pitch controls to track pilot commands while also minimizing pitch link loads. The nonlinear simulation is developed in FLIGHTLAB® with structural models of the rotor blades and control system. This model must be linearized to a linear time-invariant (LTI) system to support linear Dynamic Inversion control design. The vehicle dynamics and critical fatigue load are modeled with a linear time-periodic (LTP) model which is converted via harmonic decomposition into a high-order LTI model. This model is then reduced to design controllers across a range of airspeeds. The controllers are tested both in linear model simulations and using the full nonlinear FLIGHTLAB® model. The results show that the load alleviating controller achieves significant reduction in the pitch link peak-to-peak loads with minimal change in response characteristics, indicating that load alleviation can be achieved with no degradation in handling qualities.