... airspeed indicator instrument calibration error, may not exceed three percent or five knots, whichever is... of a steady climbing condition. (g) The effects of airspeed indicating system lag may not introduce... equivalent means of preventing malfunction due to icing. (j) Where duplicate airspeed indicators are...
Hansen, Søren; Blanke, Mogens
Airspeed sensor faults are common causes for incidents with unmanned aerial vehicles with pitot tube clogging or icing being the most common causes. Timely diagnosis of such faults or other artifacts in signals from airspeed sensing systems could potentially prevent crashes. This paper employs...
National Aeronautics and Space Administration — Micron-thin surface hot-film signatures will be used to simultaneously obtain airspeed and flow direction. The flow-angle and airspeed sensor system (FASS) will...
J.D. McLaren; J. Shamoun; C.J. Camphuysen; W. Bouten
Birds in flight are proposed to adjust their body orientation (heading) and airspeed to wind conditions adaptively according to time and energy constraints. Airspeeds in goal-directed flight are predicted to approach or exceed maximum-range airspeeds, which minimize transport costs (energy expenditu
National Aeronautics and Space Administration — Micron-thin surface hot-film gages are used to develop flow-angle and airspeed sensor system (FASS). Unlike Pitot-static and other pressure-based devices, which...
W. A. Cooper
Full Text Available A new laser air-motion sensor measures the true airspeed with a standard uncertainty of less than 0.1 m s−1 and so reduces uncertainty in the measured component of the relative wind along the longitudinal axis of the aircraft to about the same level. The calculated pressure expected from that airspeed at the inlet of a pitot tube then provides a basis for calibrating the measurements of dynamic and static pressure, reducing standard uncertainty in those measurements to less than 0.3 hPa and the precision applicable to steady flight conditions to about 0.1 hPa. These improved measurements of pressure, combined with high-resolution measurements of geometric altitude from the global positioning system, then indicate (via integrations of the hydrostatic equation during climbs and descents that the offset and uncertainty in temperature measurement for one research aircraft are +0.3 ± 0.3 °C. For airspeed, pressure and temperature, these are significant reductions in uncertainty vs. those obtained from calibrations using standard techniques. Finally, it is shown that although the initial calibration of the measured static and dynamic pressures requires a measured temperature, once calibrated these measured pressures and the measurement of airspeed from the new laser air-motion sensor provide a measurement of temperature that does not depend on any other temperature sensor.
The aerial electrostatic spraying system patented by the USDA ARS is a unique aerial application system which inductively charges spray particles for the purpose of increasing deposition and efficacy. While this system has many potential benefits, very little is known about how changes in airspeed o...
Copp, Martin R; Fetner, Mary W
Time-history data of airspeed, altitude, and acceleration obtained with the NACA VGH recorder from a twin-engine airplane operated by a regional feeder airline in the Rocky Mountains are evaluated to determine the magnitude and frequency of occurrence of gusts and gust accelerations and the operating airspeeds and altitudes. The results obtained are compared with the results previously obtained from a representative short-haul and long-haul operation.
Fixed wing Unmanned Aerial Vehicles (UAVs) are an increasingly common sensing platform, owing to their key advantages: speed, endurance and ability to explore remote areas. While these platforms are highly efficient, they cannot easily be equipped with air data sensors commonly found on their larger scale manned counterparts. Indeed, such sensors are bulky, expensive and severely reduce the payload capability of the UAVs. In consequence, UAV controllers (humans or autopilots) have little information on the actual mode of operation of the wing (normal, stalled, spin) which can cause catastrophic losses of control when flying in turbulent weather conditions. In this article, we propose a real-time air parameter estimation scheme that can run on commercial, low power autopilots in real-time. The computational method is based on a hybrid decomposition of the modes of operation of the UAV. A Bayesian approach is considered for estimation, in which the estimated airspeed, angle of attack and sideslip are described statistically. An implementation on a UAV is presented, and the performance and computational efficiency of this method are validated using hardware in the loop (HIL) simulation and experimental flight data and compared with classical Extended Kalman Filter estimation. Our benchmark tests shows that this method is faster than EKF by up to two orders of magnitude. © 2015 IEEE.
Highlights: ► Case study demonstrates application of graphical selection methodology for preliminary design. ► Small wind turbine blade example used to challenge method. ► Many designs considered, varying composite material and layup. ► Structural finite element analysis used to assess performance. ► Graphical stages used to interrogate database of solutions to select appropriate designs. -- Abstract: A small low air-speed wind turbine blade case study is used to demonstrate the effectiveness of a materials and design selection methodology described by Monroy Aceves et al. (2008) for composite structures. The blade structure comprises a shell of uniform thickness and a unidirectional reinforcement. The shell outer geometry is fixed by aerodynamic considerations. A wide range of lay-ups are considered for the shell and reinforcement. Structural analysis is undertaken using the finite element method. Results are incorporated into a database for analysis using material selection software. A graphical selection stage is used to identify the lightest blade meeting appropriate design constraints. The proposed solution satisfies the design requirements and improves on the prototype benchmark by reducing the mass by almost 50%. The flexibility of the selection software in allowing identification of trends in the results and modifications to the selection criteria is demonstrated. Introducing a safety factor of two on the material failure stresses increases the mass by only 11%. The case study demonstrates that the proposed design methodology is useful in preliminary design where a very wide range of cases should be considered using relatively simple analysis.
李超; 严家明; 刘松林
设计了以ARM微处理器为核心的中央控制处理单元,用于无人机真空速测量系统中.通过公式分解,并采用线性低次插值算法,有效解决了真空速解算公式复杂的缺点；同时利用ARM的UART串口总线,实现了传感器输出特性曲线、真空速的上位机实时显示,使系统相对误差控制在2.5％以内.测试结果证明了该系统具有良好的实效性和稳定性、精度高,优于传统的测量装置.%A central control process unit with ARM as its core is designed in order to implement the airspeed measuring system for the ummanned aerial vehicle. By the formula decomposition, and linear tow-interpolation algorithm, an effective solution is made to the shortcomings of the airspeed solution formula for calculating complex. At the same time, taking advantage of the ARM UART serial bus, the sensor output characteristic curve and real airspeed real time showed in host computer are achieved, so that the relative error of measuring system less is within 2.5%. The test results prove that the system has good effctiveness,stability and high accuracy,better than the traditional system.
In order to improve the identification accuracy of mortar trajectory,on the basis of the mortar shell centroid motion differential equation,a trajectory identification method is proposed based on air-speed sensor.By measuring velocity of projectile and changes of dynamic and static pressure,muzzle ve-locity and firing angle are identified separately,thus solving the problem of prolonging calculating time with usual method which reduces interval identification velocity and firing angle through continual itera-tion.Through analysis of ballistic simulation data,the feasibility of the algorithm is proved theoretically. Simulation result shows that trajectory identification accuracy of the algorithm is high;it is suitable for mortar projectile trajectory identification.%为提高迫击炮弹弹道辨识精度，在迫击炮弹质心运动微分方程的基础上，提出了基于空速传感器的弹道辨识方法。通过测量弹丸速度及所受动压和静压的变化，来相互独立进行初速和射角的辨识，从而解决了以往方法中通过不断迭代来缩小辨识初速和射角的区间而延长了解算时间的问题。通过对弹道仿真数据的分析，从理论上证明了该算法的可行性。仿真结果表明：该方法对弹道的识别精度较高，适合于迫击炮弹的弹道辨识。
... variations (such as horizontal gusts), and the penetration of jet streams or cold fronts), instrument errors, airframe production variations, and must not be less than Mach 0.05. (c) Design maneuvering speed V A. For... coefficients, C NA ; and (ii) n is the limit maneuvering load factor used in design (2) The value of V A...
...) The minimum speed margin must be enough to provide for atmospheric variations (such as horizontal... variations. These factors may be considered on a probability basis. The margin at altitude where MC is...); g=acceleration due to gravity (ft/sec2); a=slope of the airplane normal force coefficient curve,...
With a number of new spray testing laboratories going into operation within the U.S. and each gearing up to measure spray atomization from agricultural spray nozzles using laser diffraction, establishing and following a set of scientific standard procedures is crucial to long term data generation an...
The aerial electrostatic spraying system patented by the USDA-ARS is a unique aerial application system which inductively charges spray droplets for the purpose of increasing deposition and efficacy. While this system has many potential benefits, no published data exits which describe how changes i...
... airspeed mismatch between the pilot and co-pilot's airspeed indicators, which occurred during or after... AD, the regulatory evaluation, any comments received, and other information. The street address for... ``the MCAI''), to correct an unsafe condition for the specified products. The MCAI states: A number...
Vaughn, C. R.
NASA contribution to radar entomology is presented. Wallops Flight Center is described in terms of its radar systems. Radar tracking of birds and insects was recorded from helicopters for airspeed and vertical speed.
National Aeronautics and Space Administration — Airfoils produce more lift and less drag when the boundary layer is attached to the airfoil. With most aircraft there are combinations of airspeed and angle of...
Kaehler. Theodore J.
The Naval Aviation Enterprise (NAE) has created a program called AIRSpeed to deliver the efficiency gains of continuous process improvement (CPI). NAE leadership seeks a self-assessment tool to measure how well AIRSpeed has been implemented, including possible areas for improvement. This thesis studies the origins of continuous process improvement, the value of assessment, and current assessment methodologies. Key concepts are cited for the use of organizational assessment tools. The objectiv...
Hedenström, Anders; Åkesson, Susanne
Flight is an economical mode of locomotion, because it is both fast and relatively cheap per unit of distance, enabling birds to migrate long distances and obtain food over large areas. The power required to fly follows a U-shaped function in relation to airspeed, from which context dependent 'optimal' flight speeds can be derived. Crosswinds will displace birds away from their intended track unless they make compensatory adjustments of heading and airspeed. We report on flight track measurements in five geometrically similar tern species ranging one magnitude in body mass, from both migration and the breeding season at the island of Öland in the Baltic Sea. When leaving the southern point of Öland, migrating Arctic and common terns made a 60° shift in track direction, probably guided by a distant landmark. Terns adjusted both airspeed and heading in relation to tail and side wind, where coastlines facilitated compensation. Airspeed also depended on ecological context (searching versus not searching for food), and it increased with flock size. Species-specific maximum range speed agreed with predicted speeds from a new aerodynamic theory. Our study shows that the selection of airspeed is a behavioural trait that depended on a complex blend of internal and external factors.This article is part of the themed issue 'Moving in a moving medium: new perspectives on flight'.
Hedenström, Anders; Åkesson, Susanne
Flight is an economical mode of locomotion, because it is both fast and relatively cheap per unit of distance, enabling birds to migrate long distances and obtain food over large areas. The power required to fly follows a U-shaped function in relation to airspeed, from which context dependent 'optimal' flight speeds can be derived. Crosswinds will displace birds away from their intended track unless they make compensatory adjustments of heading and airspeed. We report on flight track measurements in five geometrically similar tern species ranging one magnitude in body mass, from both migration and the breeding season at the island of Öland in the Baltic Sea. When leaving the southern point of Öland, migrating Arctic and common terns made a 60° shift in track direction, probably guided by a distant landmark. Terns adjusted both airspeed and heading in relation to tail and side wind, where coastlines facilitated compensation. Airspeed also depended on ecological context (searching versus not searching for food), and it increased with flock size. Species-specific maximum range speed agreed with predicted speeds from a new aerodynamic theory. Our study shows that the selection of airspeed is a behavioural trait that depended on a complex blend of internal and external factors.This article is part of the themed issue 'Moving in a moving medium: new perspectives on flight'. PMID:27528786
Franklin, J. A.; Hynes, C. S.
Experiments were conducted on simulators and on the Quiet Short-Haul Research Aircraft to evaluate the effect of highly augmented control modes and electronic displays on the ability of pilots to execute precision approaches and landings on a short runway. It is found that the primary benefits of highly augmented flightpath and airspeed controls and electronic displays are realized when the pilot is required to execute precisely a complex transition and approach under instrument conditions and in the presence of a wide range of wind and turbulence conditions. A flightpath and airspeed command and stabilization system incorporating nonlinear, inverse system concepts produced fully satisfactory flightpath control throughout the aircraft's terminal operating envelope.
... recording intervals specified in appendix D of this part: (1) Time; (2) Altitude; (3) Airspeed; (4) Vertical...) Vertical acceleration; (5) Heading; (6) Time of each radio transmission either to or from air traffic... airplane equipped with a digital data bus and ARINC 717 digital flight data acquisition unit (DFDAU)...
Foster, John V. (Inventor); Cunningham, Kevin (Inventor)
A GPS-based pitot-static calibration system uses global output-error optimization. High data rate measurements of static and total pressure, ambient air conditions, and GPS-based ground speed measurements are used to compute pitot-static pressure errors over a range of airspeed. System identification methods rapidly compute optimal pressure error models with defined confidence intervals.
... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Minimum control speed. 23.149 Section 23... Maneuverability § 23.149 Minimum control speed. (a) VMC is the calibrated airspeed at which, when the critical... still inoperative, and thereafter maintain straight flight at the same speed with an angle of bank...
... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Stall speed. 25.103 Section 25.103... STANDARDS: TRANSPORT CATEGORY AIRPLANES Flight Performance § 25.103 Stall speed. (a) The reference stall speed, VSR, is a calibrated airspeed defined by the applicant. VSR may not be less than a 1-g...
... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Maximum operating limit speed. 25.1505... Operating Limitations § 25.1505 Maximum operating limit speed. The maximum operating limit speed (V MO/M MO airspeed or Mach Number, whichever is critical at a particular altitude) is a speed that may not...
... overspeed warning computer, pilot and copilot airspeed indicators, Vne converter, and AFCS air data computer... replacing certain components of the air data system. This AD was prompted by the discovery of incorrect....gov . SUPPLEMENTARY INFORMATION: Discussion On October 22, 2012, at 77 FR 64439, the Federal...
An effort is underway to update the USDA ARS aerial spray nozzle models using new droplet sizing instrumen-tation and measurement techniques. As part of this effort, the applicable maximum airspeed is being increased from 72 to 80 m/s to provide guidance to applicators when using new high speed air...
Spray adjuvants can have a substantial impact on spray atomization from agricultural nozzles; however, this process is also affected by the nozzle type, operating pressure and, for aerial application, the airspeed of application. Different types of ground spray nozzle can dramatically affect the im...
... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane instruments and equipment. 121.303... Airplane instruments and equipment. (a) Unless otherwise specified, the instrument and equipment... airspeed limitation and item of related information in the Airplane Flight Manual and pertinent...
... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Flight and navigational equipment. 121.305... Flight and navigational equipment. No person may operate an airplane unless it is equipped with the following flight and navigational instruments and equipment: (a) An airspeed indicating system with...
... Indicated Airspeed ICAO International Civil Aviation Organization ICATEE International Committee for.... 106(f), which vests final authority in the Administrator for carrying out all functions, powers, and... proposed rulemaking (NPRM) published in the Federal Register on January 12, 2009 (74 FR 1280). List...
The current USDA ARS Aerial Spray Nozzle Models were updated to reflect both new standardized measurement methods and systems, as well as, to increase operational spray pressure, aircraft airspeed and nozzle orientation angle limits. The new models were developed using both Central Composite Design...
Crowley, J W , Jr
These experiments were made at the request of the Bureau of Aeronautics, Navy Department, to investigate the velocity of the air in the slipstream in horizontal and climbing flight to determine the form of expression giving the slipstream velocity in terms of the airspeed of the airplane. The method used consisted in flying the airplane both on a level course and in climb at full throttle and measuring the slipstream velocity at seven points in the slipstream for the whole speed range of the airplane in both conditions. In general the results show that for both condition, horizontal and climbing flights, the slipstream velocity v subscript 3 and airspeed v can be represented by straight lines and consequently the equations are of the form: v subscript s = mv+b where m and b are constant. (author)
Wildmann, N.; Ravi, S.; Bange, J.
This study deals with the problem of turbulence measurement with small remotely piloted aircraft (RPA). It shows how multi-hole probes (MHPs) can be used to measure fluctuating parts of the airflow in flight up to 20 Hz. Accurate measurement of the transient wind in the outdoor environment is needed for the estimation of the 3-D wind vector as well as turbulent fluxes of heat, momentum, water vapour, etc. In comparison to an established MHP system, experiments were done to show how developments of the system setup can improve data quality. The study includes a re-evaluation of the pneumatic tubing setup, the conversion from pressures to airspeed, the pressure transducers, and the data acquisition system. In each of these fields, the steps that were taken lead to significant improvements. A spectral analysis of airspeed data obtained in flight tests shows the capability of the system to measure atmospheric turbulence up to the desired frequency range.
Abbot, K. H.; Knox, C. E.
Descent guidance was developed to provide a pilot with information to ake a fuel-conservative descent and cross a designated geographical waypoint at a preselected altitude and airspeed. The guidance was designed to reduce fuel usage during the descent and reduce the mental work load associated with planning a fuel-conservative descent. A piloted simulation was conducted to evaluate the operational use of this guidance concept. The results of the simulation tests show that the use of the guidance reduced fuel consumption and mental work load during the descent. Use of the guidance also decreased the airspeed error, but had no effect on the altitude error when the designated waypoint was crossed. Physical work load increased with the use of the guidance, but remained well within acceptable levels. The pilots found the guidance easy to use as presented and reported that it would be useful in an operational environment.
Lie, F. Adhika Pradipta
A method for estimating airspeed, angle of attack, and sideslip without using conventional, pitot-static airdata system is presented. The method relies on measurements from GPS, an inertial measurement unit (IMU) and a low-fidelity model of the aircraft's dynamics which are fused using two, cascaded Extended Kalman Filters. In the cascaded architecture, the first filter uses information from the IMU and GPS to estimate the aircraft's absolute velocity and attitude. These estimates are used as the measurement updates for the second filter where they are fused with the aircraft dynamics model to generate estimates of airspeed, angle of attack and sideslip. Methods for dealing with the time and inter-state correlation in the measurements coming from the first filter are discussed. Simulation and flight test results of the method are presented. Simulation results using high fidelity nonlinear model show that airspeed, angle of attack, and sideslip angle estimation errors are less than 0.5 m/s, 0.1 deg, and 0.2 deg RMS, respectively. Factors that affect the accuracy including the implication and impact of using a low fidelity aircraft model are discussed. It is shown using flight tests that a single linearized aircraft model can be used in lieu of a high-fidelity, non-linear model to provide reasonably accurate estimates of airspeed (less than 2 m/s error), angle of attack (less than 3 deg error), and sideslip angle (less than 5 deg error). This performance is shown to be relatively insensitive to off-trim attitudes but very sensitive to off-trim velocity.
Small Unmanned Air Vehicles (UAVs) are versatile tools with both civilian and military applications. Fixed wing UAVs require forward airspeed to remain airborne, usually resulting in constant energy expenditure to loiter over targets. A UAV capable of perching could reduce energy expenditure by settling on a site near the target, thus increasing mission duration. Avian perching techniques were observed to build a hypothesis for the biological control techniques employed during the landing man...
Flight control laws of modern aircraft are scheduled with respect to flight point parameters. The loss of the air data measurement system implies inevitably the loss of relevant scheduling information. A strategy to design a fault tolerant longitudinal flight control system is proposed which can accommodate the total loss of the angle of attack and the calibrated airspeed measurements. In this scenario the described robust longitudinal control law is employed ensuring a control performance ...
Claudel, Christian G.
A method, system, and sensor for air flow sensing. The system can include a cantilever, a transducer, and a processing module. The method can include measuring beam deflections of one or more cantilevers, extracting information about air flow, and determining one or more of an airspeed, an angle of attack, and a sideslip, based on extracted information. The system and method can exploit nonlinearities in the behavior of the cantilever in fluid flow.
Itasse, Maxime; Moschetta, Jean-Marc; Ameho, Yann; Carr, Ryan
Wind tunnel testing was performed on a VTOL aircraft in order to characterize longitudinal flight behavior during an equilibrium transition between vertical and horizontal flight modes. Trim values for airspeed, pitch, motor speed and elevator position were determined. Data was collected by independently varying the trim parameters, and stability and control derivatives were identified as functions of the trim pitch angle. A linear fractional representation model was then proposed, along with...
Taylor, Graham K; Reynolds, Kate V; Thomas, Adrian L R
Here, we analyse the energetics, performance and optimization of flight in a moving atmosphere. We begin by deriving a succinct expression describing all of the mechanical energy flows associated with gliding, dynamic soaring and thermal soaring, which we use to explore the optimization of gliding in an arbitrary wind. We use this optimization to revisit the classical theory of the glide polar, which we expand upon in two significant ways. First, we compare the predictions of the glide polar for different species under the various published models. Second, we derive a glide optimization chart that maps every combination of headwind and updraft speed to the unique combination of airspeed and inertial sink rate at which the aerodynamic cost of transport is expected to be minimized. With these theoretical tools in hand, we test their predictions using empirical data collected from a captive steppe eagle (Aquila nipalensis) carrying an inertial measurement unit, global positioning system, barometer and pitot tube. We show that the bird adjusts airspeed in relation to headwind speed as expected if it were seeking to minimize its aerodynamic cost of transport, but find only weak evidence to suggest that it adjusts airspeed similarly in response to updrafts during straight and interthermal glides.This article is part of the themed issue 'Moving in a moving medium: new perspectives on flight'. PMID:27528788
Ortega-Jimenez, Victor M; Sapir, Nir; Wolf, Marta; Variano, Evan A; Dudley, Robert
Animal fliers frequently move through a variety of perturbed flows during their daily aerial routines. However, the extent to which these perturbations influence flight control and energetic expenditure is essentially unknown. Here, we evaluate the kinematic and metabolic consequences of flight within variably sized vortex shedding flows using five Anna's hummingbirds feeding from an artificial flower in steady control flow and within vortex wakes produced behind vertical cylinders. Tests were conducted at three horizontal airspeeds (3, 6 and 9 m s(-1)) and using three different wake-generating cylinders (with diameters equal to 38, 77 and 173% of birds' wing length). Only minimal effects on wing and body kinematics were demonstrated for flight behind the smallest cylinder, whereas flight behind the medium-sized cylinder resulted in significant increases in the variances of wingbeat frequency, and variances of body orientation, especially at higher airspeeds. Metabolic rate was, however, unchanged relative to that of unperturbed flight. Hummingbirds flying within the vortex street behind the largest cylinder exhibited highest increases in variances of wingbeat frequency, and of body roll, pitch and yaw amplitudes at all measured airspeeds. Impressively, metabolic rate under this last condition increased by up to 25% compared with control flights. Cylinder wakes sufficiently large to interact with both wings can thus strongly affect stability in flight, eliciting compensatory kinematic changes with a consequent increase in flight metabolic costs. Our findings suggest that vortical flows frequently encountered by aerial taxa in diverse environments may impose substantial energetic costs.
Bäckman, J; Alerstam, T
Swifts, Apus apus, spend the night aloft and this offers an opportunity to test the degree of adaptability of bird orientation and flight to different ecological situations. We predicted the swifts' behaviour by assuming that they are adapted to minimize energy expenditure during the nocturnal flight and during a compensatory homing flight if they become displaced by wind. We tested the predictions by recording the swifts' altitudes, speeds and directions under different wind conditions with tracking radar; we found an agreement between predictions and observations for orientation behaviour, but not for altitude and speed regulation. The swifts orientated consistently into the head wind, with angular concentration increasing with increasing wind speed. However, contrary to our predictions, they did not select altitudes with slow or moderate winds, nor did they increase their airspeed distinctly when flying into strong head winds. A possible explanation is that their head-wind orientation is sufficient to keep nocturnal displacement from their home area within tolerable limits, leaving flight altitude to be determined by other factors (correlated with temperature), and airspeed to show only a marginal increase in strong winds. The swifts were often moving "backwards", heading straight into the wind but being overpowered by wind speeds exceeding their airspeed. The regular occurrence of such flights is probably uniquely associated with the swifts' remarkable habit of roosting on the wing.
Sapir, Nir; Horvitz, Nir; Dechmann, Dina K N; Fahr, Jakob; Wikelski, Martin
When animals move, their tracks may be strongly influenced by the motion of air or water, and this may affect the speed, energetics and prospects of the journey. Flying organisms, such as bats, may thus benefit from modifying their flight in response to the wind vector. Yet, practical difficulties have so far limited the understanding of this response for free-ranging bats. We tracked nine straw-coloured fruit bats (Eidolon helvum) that flew 42.5 ± 17.5 km (mean ± s.d.) to and from their roost near Accra, Ghana. Following detailed atmospheric simulations, we found that bats compensated for wind drift, as predicted under constant winds, and decreased their airspeed in response to tailwind assistance such that their groundspeed remained nearly constant. In addition, bats increased their airspeed with increasing crosswind speed. Overall, bats modulated their airspeed in relation to wind speed at different wind directions in a manner predicted by a two-dimensional optimal movement model. We conclude that sophisticated behavioural mechanisms to minimize the cost of transport under various wind conditions have evolved in bats. The bats' response to the wind is similar to that reported for migratory birds and insects, suggesting convergent evolution of flight behaviours in volant organisms.
Clark, Christopher James; Dudley, Robert
Aerodynamic theory predicts that the mechanical costs of flight are lowest at intermediate flight speeds; metabolic costs of flight should trend similarly if muscle efficiency is constant. We measured metabolic rates for nine Anna's hummingbirds (Calypte anna) and two male Allen's hummingbirds (Selasphorus sasin) feeding during flight from a free-standing mask over a range of airspeeds. Ten of 11 birds exhibited higher metabolic costs during hovering than during flight at intermediate airspeeds, whereas one individual exhibited comparable costs at hovering and during forward flight up to speeds of approximately 7 m s(-1). Flight costs of all hummingbirds increased at higher airspeeds. Relative to Anna's hummingbirds, Allen's hummingbirds exhibited deeper minima in the power curve, possibly due to higher wing loadings and greater associated costs of induced drag. Although feeding at a mask in an airstream may reduce body drag and, thus, the contributions of parasite power to overall metabolic expenditure, these results suggest that hummingbird power curves are characterized by energetic minima at intermediate speeds relative to hovering costs. PMID:20455711
Marek, C. J.; Olsen, W. A., Jr.
To correctly simulate flight in natural icing conditions, the turbulence in an icing simulator must be as low as possible. But some turbulence is required to mix the droplets from the spray nozzles and achieve an icing cloud of uniform liquid water content. The goal for any spray system is to obtain the widest possible spray cloud with the lowest possible turbulence in the test section of a icing tunnel. This investigation reports the measurement of turbulence and the three-dimensional spread of the cloud from a single spray nozzle. The task was to determine how the air turbulence and cloud width are affected by spray bars of quite different drag coefficients, by changes in the turbulence upstream of the spray, the droplet size, and the atomizing air. An ice accretion grid, located 6.3 m downstream of the single spray nozzle, was used to measure cloud spread. Both the spray bar and the grid were located in the constant velocity test section. Three spray bar shapes were tested: the short blunt spray bar used in the NASA Lewis Icing Research Tunnel, a thin 14.6 cm chord airfoil, and a 53 cm chord NACA 0012 airfoil. At the low airspeed (56 km/hr) the ice accretion pattern was axisymmetric and was not affected by the shape of the spray bar. At the high airspeed (169 km/hr) the spread was 30 percent smaller than at the low airspeed. For the widest cloud the spray bars should be located as far upstream in the low velocity plenum of the icing tunnel. Good comparison is obtained between the cloud spread data and predicitons from a two-dimensional cloud mixing computer code using the two equation turbulence (k epsilon g) model.
Sapir, Nir; Dudley, Robert
Backward flight is a frequently used transient flight behavior among members of the species-rich hummingbird family (Trochilidae) when retreating from flowers, and is known from a variety of other avian and hexapod taxa, but the biomechanics of this intriguing locomotor mode have not been described. We measured rates of oxygen uptake (V(O2)) and flight kinematics of Anna's hummingbirds, Calypte anna (Lesson), within a wind tunnel using mask respirometry and high-speed videography, respectively, during backward, forward and hovering flight. We unexpectedly found that in sustained backward flight is similar to that in forward flight at equivalent airspeed, and is about 20% lower than hovering V(O2). For a bird that was measured throughout a range of backward airspeeds up to a speed of 4.5 m s(-1), the power curve resembled that of forward flight at equivalent airspeeds. Backward flight was facilitated by steep body angles coupled with substantial head flexion, and was also characterized by a higher wingbeat frequency, a flat stroke plane angle relative to horizontal, a high stroke plane angle relative to the longitudinal body axis, a high ratio of maximum:minimum wing positional angle, and a high upstroke:downstroke duration ratio. Because of the convergent evolution of hummingbird and some hexapod flight styles, flying insects may employ similar kinematics while engaged in backward flight, for example during station keeping or load lifting. We propose that backward flight behavior in retreat from flowers, together with other anatomical, physiological, morphological and behavioral adaptations, enables hummingbirds to maintain strictly aerial nectarivory. PMID:23014570
Kohlman, D. L.; Hammer, J.
Developments in aerodyamic, structural and propulsion technologies which influence the potential for significant improvements in performance and fuel efficiency of general aviation business airplanes are discussed. The advancements include such technolgies as natural laminar flow, composite materials, and advanced intermittent combustion engines. The design goal for this parameter design study is a range of 1300 nm at 300 knots true airspeed with a payload of 1200lbs at 35,000 ft cruise altitude. The individual and synergistic effects of various advanced technologies on the optimization of this class of high performance, single engine, propeller driven business airplanes are identified.
Steen, Laura E.; VanZante, Judith Foss; Broeren, Andy P.; Kubiak, Mark J.
In 2011, the heat exchanger and refrigeration plant for NASA Glenn Research Centers Icing Research Tunnel (IRT) were upgraded. Flow quality surveys were performed in the settling chamber of the IRT in order to understand the effect that the new heat exchanger had on the flow quality upstream of the spray bars. Measurements were made of the total pressure, static pressure, total temperature, airspeed, and flow angle (pitch and yaw). These measurements were directly compared to measurements taken in 2000, after the previous heat exchanger was installed. In general, the flow quality appears to have improved with the new heat exchanger.
Hansen, Søren; Blanke, Mogens; Adrian, Jens
Unmanned Aerial Vehicles need a large degree of tolerance to faults. One of the most important steps towards this is the ability to detect and isolate faults in sensors and actuators in real time and make remedial actions to avoid that faults develop to failure. This paper analyses...... the possibilities of detecting faults in the pitot tube of a small unmanned aerial vehicle, a fault that easily causes a crash if not diagnosed and handled in time. Using as redundant information the velocity measured from an onboard GPS receiver, the air-speed estimated from engine throttle and the pitot tube...
The load capacity of airplane's ram air turbine ( RAT) depends not only on airplane airspeed, but al-so on the flight attitude. It is difficult to calculate by using theoretical equations. A flight test method is designed to determine the RAT's real-time load capacity by just analyzing some flight test data obtained. This method plans two test subjects:RAT load capacity test at level flight and RAT load capacity test at sideslip. The former test can obtain the variation of RAT airspeed with airplane airspeed for different weights, and the latter can bring us the re-lationship for different sideslip angles. The results of the two tests show that the RAT airspeed at any weight and flight attitude can be determined and further know its load capacity.%飞机RAT的实际带载能力不仅取决于飞机空度，而且受到飞行姿态的制约，很难用理论公式准确推算。提出了一种飞行试验方法，分析特定的试飞测试数据，可获得飞机RAT的实际带载能力。试验包括RAT平飞带载和RAT侧滑带载两部分试飞。通过RAT平飞带载试验，得到不同重量下的RAT空速与飞机空速的关系曲线；通过RAT侧滑带载试验，可得到不同侧滑角飞行下RAT空速与飞机空速的关系曲线。结合试验结果，可以确定飞机在任意重量和飞行姿态下的RAT空速，进而获知RAT实际带载能力。
We extend the use of cantilever beams as flow sensors for small aircraft. As such, we propose a novel method to measure the airspeed and the angle of attack at which the air travels across a small flying vehicle. We measure beam deflections and extract information about the surrounding flow. Thus, we couple a nonlinear beam model with a potential flow simulator through a fluid-structure interaction scheme. We use this numerical approach to generate calibration curves that exhibit the trend for the variations of the limit cycle oscillations amplitudes of flexural and torsional vibrations with the air speed and the angle of attack, respectively. © The Author(s) 2013.
Liang Kun; Li Rong; Chen Jianjun; Ma Jiantai
The catalysts of CeO2 and the mixture of CeO2 and CuO were prepared, and the activities of these catalysts for completely oxidizing benzene were studied.The results show that the optimal proportion of CeO2/CuO is 6: 4.The highest temperature at which benzene was completely oxidized on these catalysts at different airspeed was measured.Compared these catalysts with the noble metal used, our catalysts had superiority in the resources and the industrial cost besides good activities.
Pines, S.; Schmidt, S. F.
The results of a study to determine scaling, storage, and word length requirements for programming the Kalman filter on the GE-701 Whole Word Computer are reported. Simulation tests are presented which indicate that the Kalman filter, using a square root formulation with process noise added, utilizing MLS, radar altimeters, and airspeed as navigation aids, may be programmed for the GE-701 computer to successfully navigate and control the Boeing B737-100 during landing approach, landing rollout, and turnoff. The report contains flow charts, equations, computer storage, scaling, and word length recommendations for the Kalman filter on the GE-701 Whole Word computer.
Straub, Jeremy; Huber, Justin
The validation of safety-critical applications, such as autonomous UAV operations in an environment which may include human actors, is an ill posed problem. To confidence in the autonomous control technology, numerous scenarios must be considered. This paper expands upon previous work, related to autonomous testing of robotic control algorithms in a two dimensional plane, to evaluate the suitability of similar techniques for validating artificial intelligence control in three dimensions, where a minimum level of airspeed must be maintained. The results of human-conducted testing are compared to this automated testing, in terms of error detection, speed and testing cost.
Here we report a robust thermal anemometer which can be easily built. It was conceived to measure outdoor wind speeds, and for airspeed monitoring in wind tunnels and other indoor uses. It works at a constant, low temperature of approximately 90$^\\circ$C, so that an independent measurement of the air temperature is required to give a correct speed reading. Despite the size and high thermal inertia of the probe, the test results show that this anemometer is capable of measuring turbulent fluctuations up to ~100 Hz in winds of ~14 m/s, which corresponds to a scale similar to the length of the probe.
Al-Fifi, Salman Amsari
The topic of my thesis is Experimental and Numerical Study of Open Air Active Cooling. The present research is intended to investigate experimentally and Numerically the effectiveness of cooling large open areas like stadiums, shopping malls, national gardens, amusement parks, zoos, transportation facilities and government facilities or even in buildings outdoor gardens and patios. Our cooling systems are simple cooling fans with different diameters and a mist system. This type of cooling systems has been chosen among the others to guarantee less energy consumption, which will make it the most favorable and applicable for cooling such places mentioned above. In the experiments, the main focus is to study the temperature domain as a function of different fan diameters aerodynamically similar in different heights till we come up with an empirical relationship that can determine the temperature domain for different fan diameters and for different heights of these fans. The experimental part has two stages. The first stage is devoted to investigate the maximum range of airspeed and profile for three different fan diameters and for different heights without mist, while the second stage is devoted to investigate the maximum range of temperature and profile for the three different diameter fans and for different heights with mist. The computational study is devoted to built an experimentally verified mathematical model to be used in the design and optimization of water mist cooling systems, and to compare the mathematical results to the experimental results and to get an insight of how to apply such evaporative mist cooling for different places for different conditions. In this study, numerical solution is presented based on experimental conditions, such dry bulb temperature, wet bulb temperature, relative humidity, operating pressure and fan airspeed. In the computational study, all experimental conditions are kept the same for the three fans except the fan airspeed
Sheridan, P. F.; Robinson, C.; Shaw, J.; White, F.
A math model was formulated to represent some of the aerodynamic effects of low speed, low altitude, and steeply descending flight. The formulation is intended to be consistent with the single rotor real time simulation model at NASA Ames Research Center. The effect of low speed, low altitude flight on main rotor downwash was obtained by assuming a uniform plus first harmonic inflow model and then by using wind tunnel data in the form of hub loads to solve for the inflow coefficients. The result was a set of tables for steady and first harmonic inflow coefficients as functions of ground proximity, angle of attack, and airspeed. The aerodynamics associated with steep descending flight in the vortex ring state were modeled by replacing the steady induced downwash derived from momentum theory with an experimentally derived value and by including a thrust fluctuations effect due to vortex shedding. Tables of the induced downwash and the magnitude of the thrust fluctuations were created as functions of angle of attack and airspeed.
Augere, B.; Besson, B.; Fleury, D.; Goular, D.; Planchat, C.; Valla, M.
Lidar (light detection and ranging) is a well-established measurement method for the prediction of atmospheric motions through velocity measurements. Recent advances in 1.5 μm Lidars show that the technology is mature, offers great ease of use, and is reliable and compact. A 1.5 μm airborne Lidar appears to be a good candidate for airborne in-flight measurement systems. It allows measurements remotely, outside aircraft aerodynamic disturbance, and absolute air speed (no need for calibration) with great precision in all aircraft flight domains. In the framework of the EU AIM2 project, the ONERA task has consisted of developing and testing a 1.5 μm anemometer sensor for in-flight airspeed measurements. The objective of this work is to demonstrate that the 1.5 μm Lidar sensor can increase the quality of the data acquisition procedure for aircraft flight test certification. This article presents the 1.5 μm anemometer sensor dedicated to in-flight airspeed measurements and describes the flight tests performed successfully on-board the Piaggio P180 aircraft. Lidar air data have been graphically compared to the air data provided by the aircraft flight test instrumentation (FTI) in the reference frame of the Lidar sensor head. Very good agreement of true air speed (TAS) by a fraction of ms‑1, angle of sideslip (AOS), and angle of attack (AOA) by a fraction of degree were observed.
Bidwell, Colin S.; Rigby, David L.
A flow and ice particle trajectory analysis was performed for the booster of the Honeywell AL502 engine. The analysis focused on two closely related conditions one of which produced a rollback and another which did not rollback during testing in the Propulsion Systems Lab at NASA Glenn Research Center. The flow analysis was generated using the NASA Glenn GlennHT flow solver and the particle analysis was generated using the NASA Glenn LEWICE3D v3.56 ice accretion software. The flow and particle analysis used a 3D steady flow, mixing plane approach to model the transport of flow and particles through the engine. The inflow conditions for the rollback case were: airspeed, 145 ms; static pressure, 33,373 Pa; static temperature, 253.3 K. The inflow conditions for the non-roll-back case were: airspeed, 153 ms; static pressure, 34,252 Pa; static temperature, 260.1 K. Both cases were subjected to an ice particle cloud with a median volume diameter of 24 microns, an ice water content of 2.0 gm3 and a relative humidity of 100 percent. The most significant difference between the rollback and non-rollback conditions was the inflow static temperature which was 6.8 K higher for the non-rollback case.
It has been hypothesized that a human pilot uses the same set of generic skills to control a wide variety of aircraft. If this is true, then it should be possible to construct an electronic controller which embodies this generic skill set such that it can successfully control difference airplanes without being matched to a specific airplane. In an attempt to create such a system, a fuzzy logic controller was devised to control throttle position and another to control elevator position. These two controllers were used to control flight path angle and airspeed for both a piston powered single engine airplane simulation and a business jet simulation. Overspeed protection and stall protection were incorporated in the form of expert systems supervisors. It was found that by using the artificial intelligence techniques of fuzzy logic and expert systems, a generic longitudinal controller could be successfully used on two general aviation aircraft types that have very difference characteristics. These controllers worked for both airplanes over their entire flight envelopes including configuration changes. The controllers for both airplanes were identical except for airplane specific limits (maximum allowable airspeed, throttle lever travel, etc.). The controllers also handled configuration changes without mode switching or knowledge of the current configuration. This research validated the fact that the same fuzzy logic based controller can control two very different general aviation airplanes. It also developed the basic controller architecture and specific control parameters required for such a general controller.
Baxley, Brian T.; Murdoch, Jennifer L.; Swieringa, Kurt A.; Barmore, Bryan E.; Capron, William R.; Hubbs, Clay E.; Shay, Richard F.; Abbott, Terence S.
The predicted increase in the number of commercial aircraft operations creates a need for improved operational efficiency. Two areas believed to offer increases in aircraft efficiency are optimized profile descents and dependent parallel runway operations. Using Flight deck Interval Management (FIM) software and procedures during these operations, flight crews can achieve by the runway threshold an interval assigned by air traffic control (ATC) behind the preceding aircraft that maximizes runway throughput while minimizing additional fuel consumption and pilot workload. This document describes an experiment where 24 pilots flew arrivals into the Dallas Fort-Worth terminal environment using one of three simulators at NASA?s Langley Research Center. Results indicate that pilots delivered their aircraft to the runway threshold within +/- 3.5 seconds of their assigned time interval, and reported low workload levels. In general, pilots found the FIM concept, procedures, speeds, and interface acceptable. Analysis of the time error and FIM speed changes as a function of arrival stream position suggest the spacing algorithm generates stable behavior while in the presence of continuous (wind) or impulse (offset) error. Concerns reported included multiple speed changes within a short time period, and an airspeed increase followed shortly by an airspeed decrease.
Full Text Available Dynamic soaring is a special flying technique designed to allow UAVs (unmanned aerial vehicles to extract energy from wind gradient field and enable UAVs to increase the endurance. In order to figure out the energy-extraction mechanisms in dynamic soaring, a noninertial wind relative reference frame of aircraft is built. In the noninertial frame, there is an inertial force which is created by gradient wind field. When the wind gradient (GW and the components of airspeed (vzvx are positive, inertial force (F makes positive work to the aircraft. In the meantime, an equilibrium position theory of dynamic soaring is proposed. At the equilibrium positions, the increased potential energy is greater than the wasted kinetic energy when the aircraft is flying upwards. The mechanical energy is increased in this way, and the aircraft can store energy for flight. According to the extreme value theory, contour line figures of the maximum function and the component of airspeed (vz are obtained to find out the aircraft’s lifting balance allowance in dynamic soaring. Moreover, this equilibrium position theory can also help to conduct an aircraft to acquire energy from the environment constantly.
Current developments surrounding the use of unmanned aerial vehicles have produced a need for a high quality data acquisition platform developed specifically a research environment. This work was undertaken to produce such a system that is low cost, extensible, and better supports fixed wing research through the inclusion of a custom vane based air data probe capable of measuring airspeed, angle of attack, and angle of sideslip. This was accomplished by starting with the open source Pixhawk system as the core and then modifying the device firmware and adding sensors to suit the needs of current aerospace research at OSU. An overview of each component of the system is presented, as well as a description of various firmware modifications to the stock Pixhawk system. Tests were then performed on all of the major sensors using bench testing, wind tunnel analysis, and flight maneuvers to determine the final performance of each part of the system. This research shows that all of the critical sensors on the data acquisition platform produce data acceptable for flight research. The accelerometer has been shown to have an overall tolerance of +/-0.0545 m/s², with +/-0.223 deg/s for the gyroscopic sensor, +/-1.32 hPa for the barometric sensor, +/-0.318 m/s for the airspeed sensor, +/-1.65 °C for the outside air temperature sensor, and +/-0.00115 V for the analog to digital converter. The stock calibration curve for the airspeed sensor was determined to be correct to within +/-0.5 in H2O through wind tunnel testing, and an experimental step input analysis on the flow direction vanes showed that worst case steady state error and time to damp are acceptable for the system. Power spectral density and spectral coherence analysis of flight data was used to show that the custom air data probe is capable of following the flight dynamics of a given aircraft to within a 10 percent tolerance across a range of frequencies. Finally, general performance of the system was proven using
Foster, John V.; Cunningham, Kevin
Pressure-based airspeed and altitude measurements for aircraft typically require calibration of the installed system to account for pressure sensing errors such as those due to local flow field effects. In some cases, calibration is used to meet requirements such as those specified in Federal Aviation Regulation Part 25. Several methods are used for in-flight pitot-static calibration including tower fly-by, pacer aircraft, and trailing cone methods. In the 1990 s, the introduction of satellite-based positioning systems to the civilian market enabled new inflight calibration methods based on accurate ground speed measurements provided by Global Positioning Systems (GPS). Use of GPS for airspeed calibration has many advantages such as accuracy, ease of portability (e.g. hand-held) and the flexibility of operating in airspace without the limitations of test range boundaries or ground telemetry support. The current research was motivated by the need for a rapid and statistically accurate method for in-flight calibration of pitot-static systems for remotely piloted, dynamically-scaled research aircraft. Current calibration methods were deemed not practical for this application because of confined test range size and limited flight time available for each sortie. A method was developed that uses high data rate measurements of static and total pressure, and GPSbased ground speed measurements to compute the pressure errors over a range of airspeed. The novel application of this approach is the use of system identification methods that rapidly compute optimal pressure error models with defined confidence intervals in nearreal time. This method has been demonstrated in flight tests and has shown 2- bounds of approximately 0.2 kts with an order of magnitude reduction in test time over other methods. As part of this experiment, a unique database of wind measurements was acquired concurrently with the flight experiments, for the purpose of experimental validation of the
Bousquet, Gabriel; Triantafyllou, Michael; Slotine, Jean-Jacques
Dynamic soaring is the flight technique where a glider, either avian or manmade, extracts its propulsive energy from the non-uniformity of horizontal winds. Albatrosses have been recorded to fly an impressive 5000 km/week at no energy cost of their own. In the sharp boundary layer limit, we show that the popular image, where the glider travels in a succession of half turns, is suboptimal for travel speed, airspeed, and soaring ability. Instead, we show that the strategy that maximizes the three criteria simultaneously is a succession of infinitely small arc-circles connecting transitions between the calm and windy layers. The model is consistent with the recordings of albatross flight patterns. This lowers the required wind speed for dynamic soaring by over 50% compared to previous beliefs. In the thick boundary layer limit, energetic considerations allow us to predict a minimum wind gradient necessary for sustained soaring consistent with numerical models.
Claudel, Christian G.
Systems and methods to protect the flight envelope in both manual flight and flight by a commercial autopilot are provided. A system can comprise: an inertial measurement unit (IMU); a computing device in data communication with the IMU; an application executable by the computing device comprising: logic that estimates an angle of attack; a slip angle; and a speed of an unmanned aerial vehicle (UAV) based at least in part on data received from the UAV. A method can comprise estimating, via a computing device, flight data of a UAV based at least in part on data received from an IMU; comparing the estimated flight data with measured flight data; and triggering an error indication in response to a determination that the measured flight data exceeds a predefined deviation of the estimated flight data. The estimated speed can comprise an estimated airspeed, vertical speed and/or ground velocity.
Bents, David J.
This report presents the conceptual design for a solar electric lighter-than-air, unmanned aerial vehicle, based on existing technology already reduced to practice, that could carry a 600-kg (1322-lbm) payload to altitudes up to 30 kft (9000 m), continuously maintain an airspeed up to 40 kt (21 m/sec), and remain in flight for up to 100 days. The design is based on modern nonrigid airship technology, high-strength polymer fabrics and barrier films, and previously demonstrated aerospace electrical power technology, including lightweight photovoltaics and hydrogen-air polymer electrolyte membrane (PEM) fuel cells. The vehicle concept exploits the inherent synergy between the use of hydrogen as a lifting gas and the use of hydrogen-air PEM fuel-cell technology for onboard solar energy storage. In this report, the air vehicle concept is physically characterized and its estimated performance envelope is defined
Bonafe, J. L.
From its first designed airplane, Airbus considered mandatory a help in the crew's decision-making process to initiate an escape maneuver and help to successfully realize it. All the Airbus airplanes designed since 1975 included an alpha-floor function and a speed reference control law imbedded in the speed reference system (SRS) box for A 300 and FAC and FCC for A 310, A300/600 and the A 320. Alpha-Floor function takes into account the airplane energy situation considering angle of attack and observed longitudinal situation in order to apply immediately the full power without any pilot action. Speed reference managers control airspeed and/or ground speed in order to survive a maximum in shear situation. In order to comply with the new FAA regulation: Aerospatiale and Airbus developed more efficient systems. A comparison between 1975 and a newly developed system is given. It is explained how the new system improves the situation.
Randleff, Lars Rosenberg
is detected the pilot may choose to deploy electronic countermeasures to avoid the impact of the missile. The countermeasures to choose depends on e.g. the type of missile and guidance system, distance and direction between the missile and the aircraft, an assessment of the environment hostility, aircraft...... altitude and airspeed, and the availability of countermeasures. Radar systems, guidance of missiles, and electronic countermeasures are all parts of the electronic warfare domain. A brief description of this domain is given. It contains an introduction to both systems working on-board the aircraft......During a mission over enemy territory a fighter aircraft may be engaged by ground based threats. The pilot can use different measures to avoid the aircraft from being detected by e.g. enemy radar systems. If the enemy detects the aircraft a missile may be fired to seek and destroy the aircraft...
WU Xue-Mei; TUO Xian-Guo; LI Zhe; LIU Ming-Zhe; ZHANG Jin-Zhao; DONG Xiang-Long; LI Ping-Chuan
Long-range alpha detectors (LRADs) are attracting much attention in the decommissioning of nuclear facilities because of some problems in obtaining source positions on an interior surface during pipe decommissioning.By utilizing the characteristic that LRAD detects alphas by collecting air-driving ions,this article applies a method to localize the radioactive source by ions' fluid property.By obtaining the ion travel time and the airspeed distribution in the pipe,the source position can be determined.Thus this method overcomes the ion's lack of periodic characteristics.Experimental results indicate that this method can approximately localize the source inside the pipe.The calculation results are in good agreement with the experimental results.
A rotorcraft-based unmanned aerial vehicle exhibits more complex properties compared to its full-size counterparts due to its increased sensitivity to control inputs and disturbances and higher bandwidth of its dynamics. As an aerial vehicle with vertical take-off and landing capability, the helicopter specifically poses a difficult problem of transition between forward flight and unstable hover and vice versa. The LPV control technique explicitly takes into account the change in performance due to the real-time parameter variations. The technique therefore theoretically guarantees the performance and robustness over the entire operating envelope. In this study, we investigate a new approach implementing model identification for use in the LPV control framework. The identification scheme employs recursive least square technique implemented on the LPV system represented by dynamics of helicopter during a transition. The airspeed as the scheduling of parameter trajectory is not assumed to vary slowly. The exclu...
Full Text Available This paper presents an experimental method for assessing the performances and the propulsion power of a UAV in several points based on telemetry. The points in which we make the estimations are chosen based on several criteria and the fallowing parameters are measured: airspeed, time-to-climb, altitude and the horizontal distance. With the estimated propulsion power and knowing the shaft motor power, the propeller efficiency is determined at several speed values. The shaft motor power was measured in the lab using the propeller as a break. Many flights, using the same UAV configuration, were performed before extracting flight data, in order to reduce the instrumental or statistic errors. This paper highlights both the methodology of processing the data and the validation of theoretical results.
López, J; Dormido, R; Dormido, S; Gómez, J P
The objective of this paper is the implementation and validation of a robust H ∞ controller for an UAV to track all types of manoeuvres in the presence of noisy environment. A robust inner-outer loop strategy is implemented. To design the H ∞ robust controller in the inner loop, H ∞ control methodology is used. The two controllers that conform the outer loop are designed using the H ∞ Loop Shaping technique. The reference vector used in the control architecture formed by vertical velocity, true airspeed, and heading angle, suggests a nontraditional way to pilot the aircraft. The simulation results show that the proposed control scheme works well despite the presence of noise and uncertainties, so the control system satisfies the requirements.
Jex, H. R.; Hogue, J. R.; Gelhausen, P.
An examination is conducted of the Skyship 500's dynamic response to control inputs from elevators, rudders, and throttles at zero, 25, and 40 kts indicated airspeed. Input frequency sweeps were made with pitch and turn controls at 25 and 40 kts, ranging in frequency from about 0.03 to 1.5 Hz. FFT data analysis was then applied to compute describing functions for each run. Frequency responses are noted to be very smooth, and comparisons between repeat runs indicate excellent agreement. Summary plots of the faired describing functions from each run form the core of the data presented. These data constitute a comprehensive and reliable data base on which to predicate future dynamic simulation mathematical models of small airship dynamic response.
Hess, R. A.
Quantitative Feedback Theory describes a frequency-domain technique for the design of multi-input, multi-output control systems which must meet time or frequency domain performance criteria when specified uncertainty exists in the linear description of the vehicle dynamics. This theory is applied to the design of the longitudinal flight control system for a linear model of the BO-105C rotorcraft. Uncertainty in the vehicle model is due to the variation in the vehicle dynamics over a range of airspeeds from 0-100 kts. For purposes of exposition, the vehicle description contains no rotor or actuator dynamics. The design example indicates the manner in which significant uncertainty exists in the vehicle model. The advantage of using a sequential loop closure technique to reduce the cost of feedback is demonstrated by example.
Rutan, Elbert L. (Inventor)
An orbital launch system and its method of operation use a maneuver to improve the launch condition of a booster rocket and payload. A towed launch aircraft, to which the booster rocket is mounted, is towed to a predetermined elevation and airspeed. The towed launch aircraft begins the maneuver by increasing its lift, thereby increasing the flight path angle, which increases the tension on the towline connecting the towed launch aircraft to a towing aircraft. The increased tension accelerates the towed launch aircraft and booster rocket, while decreasing the speed (and thus the kinetic energy) of the towing aircraft, while increasing kinetic energy of the towed launch aircraft and booster rocket by transferring energy from the towing aircraft. The potential energy of the towed launch aircraft and booster rocket is also increased, due to the increased lift. The booster rocket is released and ignited, completing the launch.
Hirst, E.; Kaye, P. H.; Greenaway, R. S.; Field, P.; Johnson, D. W.
Preliminary experimental results are presented from an aircraft-mounted probe designed to provide in situ data on cloud particle shape, size, and number concentration. In particular, the probe has been designed to facilitate discrimination between super-cooled water droplets and ice crystals of 1-25 μm size within mixed-phase clouds and to provide information on cloud interstitial aerosols. The probe acquires spatial light scattering data from individual particles at throughput rates of several thousand particles per second. These data are logged at 100 ms intervals to allow the distribution and number concentration of each particle type to be determined with 10 m spatial resolution at a typical airspeed of 100 m s -1. Preliminary results from flight data recorded in altocumulus castellanus, showing liquid water phase, mixed phase, and ice phase are presented to illustrate the probe's particle discrimination capabilities.
Williams, Daniel M.; Consiglio, Maria C.; Murdoch, Jennifer L.; Adams, Catherine H.
This paper provides an analysis of Flight Technical Error (FTE) from recent SATS experiments, called the Higher Volume Operations (HVO) Simulation and Flight experiments, which NASA conducted to determine pilot acceptability of the HVO concept for normal operating conditions. Reported are FTE results from simulation and flight experiment data indicating the SATS HVO concept is viable and acceptable to low-time instrument rated pilots when compared with today s system (baseline). Described is the comparative FTE analysis of lateral, vertical, and airspeed deviations from the baseline and SATS HVO experimental flight procedures. Based on FTE analysis, all evaluation subjects, low-time instrument-rated pilots, flew the HVO procedures safely and proficiently in comparison to today s system. In all cases, the results of the flight experiment validated the results of the simulation experiment and confirm the utility of the simulation platform for comparative Human in the Loop (HITL) studies of SATS HVO and Baseline operations.
Fadjar Rahino Triputra
Full Text Available Developing a nonlinear adaptive control system for a fixed-wing unmanned aerial vehicle (UAV requires a mathematical representation of the system dynamics analytically as a set of differential equations in the form of a strict-feedback systems. This paper presents a method for modeling a nonlinear flight dynamics of the fixed-wing UAV of BPPT Wulung in any conditions of the flight altitude and airspeed for the first step into designing a nonlinear adaptive controller. The model was formed into 10-DOF differential equations in the form of strict-feedback systems which separates the terms of elevator, aileron, rudder and throttle from the model. The model simulation results show the behavior of the flight dynamics of the Wulung UAV and also prove the compliance with the actual flight test results.
Lidar (light detection and ranging) is a well-established measurement method for the prediction of atmospheric motions through velocity measurements. Recent advances in 1.5 μm Lidars show that the technology is mature, offers great ease of use, and is reliable and compact. A 1.5 μm airborne Lidar appears to be a good candidate for airborne in-flight measurement systems. It allows measurements remotely, outside aircraft aerodynamic disturbance, and absolute air speed (no need for calibration) with great precision in all aircraft flight domains. In the framework of the EU AIM2 project, the ONERA task has consisted of developing and testing a 1.5 μm anemometer sensor for in-flight airspeed measurements. The objective of this work is to demonstrate that the 1.5 μm Lidar sensor can increase the quality of the data acquisition procedure for aircraft flight test certification. This article presents the 1.5 μm anemometer sensor dedicated to in-flight airspeed measurements and describes the flight tests performed successfully on-board the Piaggio P180 aircraft. Lidar air data have been graphically compared to the air data provided by the aircraft flight test instrumentation (FTI) in the reference frame of the Lidar sensor head. Very good agreement of true air speed (TAS) by a fraction of ms−1, angle of sideslip (AOS), and angle of attack (AOA) by a fraction of degree were observed. (special issue article)
宋辉; 陈欣; 李春涛
As one of the most complex phases in whole flight process, a high span-chord UAV is extremely sensitive to crosswind because of its low airspeed during landing. Through analyzing the influence of crosswind on flight performance of the sample UAV during landing phase, this essay puts forward some methods to enhance the control performance during automatic landing by optimizing the landing trajectory, controlling the airspeed and improving the longitudinal and lateral control strategy. Finally, simulation results show that the landing trajectory is reasonable , the automatic landing control strategy is valid and the control law is correct, which satisfy the demands of the automatic landing control under the crosswind conditions.%大展弦比无人机进场着陆过程中速度较低,对侧风十分敏感,是飞行过程中控制最复杂的阶段之一.通过分析侧风条件对样例无人机着陆过程的影响,提出了通过优化设计着陆轨迹线、控制空速以及完善纵横向控制策略的方法来改善侧风条件下大展弦比无人机自动着陆的控制品质.仿真结果表明,所设计的自动着陆轨迹线合理,控制策略完备,控制律品质良好,满足样例无人机在侧风条件下自动着陆控制的设计要求.
Ozdemir, Gurbuz Taha
A conventional helicopter has limits on performance at high speeds because of the limitations of main rotor, such as compressibility issues on advancing side or stall issues on retreating side. Auxiliary lift and thrust components have been suggested to improve performance of the helicopter substantially by reducing the loading on the main rotor. Such a configuration is called the compound rotorcraft. Rotor speed can also be varied to improve helicopter performance. In addition to improved performance, compound rotorcraft and variable RPM can provide a much larger degree of control redundancy. This additional redundancy gives the opportunity to further enhance performance and handling qualities. A flight control system is designed to perform in-flight optimization of redundant control effectors on a compound rotorcraft in order to minimize power required and extend range. This "Fly to Optimal" (FTO) control law is tested in simulation using the GENHEL model. A model of the UH-60, a compound version of the UH-60A with lifting wing and vectored thrust ducted propeller (VTDP), and a generic compound version of the UH-60A with lifting wing and propeller were developed and tested in simulation. A model following dynamic inversion controller is implemented for inner loop control of roll, pitch, yaw, heave, and rotor RPM. An outer loop controller regulates airspeed and flight path during optimization. A Golden Section search method was used to find optimal rotor RPM on a conventional helicopter, where the single redundant control effector is rotor RPM. The FTO builds off of the Adaptive Performance Optimization (APO) method of Gilyard by performing low frequency sweeps on a redundant control for a fixed wing aircraft. A method based on the APO method was used to optimize trim on a compound rotorcraft with several redundant control effectors. The controller can be used to optimize rotor RPM and compound control effectors through flight test or simulations in order to
Farnsworth, Andrew; Van DOREN, Benjamin M; Hochachka, Wesley M; Sheldon, Daniel; Winner, Kevin; Irvine, Jed; Geevarghese, Jeffrey; Kelling, Steve
Billions of birds migrate at night over North America each year. However, few studies have described the phenology of these movements, such as magnitudes, directions, and speeds, for more than one migration season and at regional scales. In this study, we characterize density, direction, and speed of nocturnally migrating birds using data from 13 weather surveillance radars in the autumns of 2010 and 2011 in the northeastern USA. After screening radar data to remove precipitation, we applied a recently developed algorithm for characterizing velocity profiles with previously developed methods to document bird migration. Many hourly radar scans contained windborne "contamination," and these scans also exhibited generally low overall reflectivities. Hourly scans dominated by birds showed nightly and seasonal patterns that differed markedly from those of low reflectivity scans. Bird migration occurred during many nights, but a smaller number of nights with large movements of birds defined regional nocturnal migration. Densities varied by date, time, and location but peaked in the second and third deciles of night during the autumn period when the most birds were migrating. Migration track (the direction to which birds moved) shifted within nights from south-southwesterly to southwesterly during the seasonal migration peaks; this shift was not consistent with a similar shift in wind direction. Migration speeds varied within nights, although not closely with wind speed. Airspeeds increased during the night; groundspeeds were highest between the second and third deciles of night, when the greatest density of birds was migrating. Airspeeds and groundspeeds increased during the fall season, although groundspeeds fluctuated considerably with prevailing winds. Significant positive correlations characterized relationships among bird densities at southern coastal radar stations and northern inland radar stations. The quantitative descriptions of broadscale nocturnal migration
Noll, Thomas E.; Ishmael, Stephen D.; Henwood, Bart; Perez-Davis, Marla E.; Tiffany, Geary C.; Madura, John; Gaier, Matthew; Brown, John M.; Wierzbanowski, Ted
The Helios Prototype was originally planned to be two separate vehicles, but because of resource limitations only one vehicle was developed to demonstrate two missions. The vehicle consisted of two configurations, one for each mission. One configuration, designated HP01, was designed to operate at extremely high altitudes using batteries and high-efficiency solar cells spread across the upper surface of its 247-foot wingspan. On August 13, 2001, the HP01 configuration reached an altitude of 96,863 feet, a world record for sustained horizontal flight by a winged aircraft. The other configuration, designated HP03, was designed for long-duration flight. The plan was to use the solar cells to power the vehicle's electric motors and subsystems during the day and to use a modified commercial hydrogen-air fuel cell system for use during the night. The aircraft design used wing dihedral, engine power, elevator control surfaces, and a stability augmentation and control system to provide aerodynamic stability and control. At about 30 minutes into the second flight of HP03, the aircraft encountered a disturbance in the way of turbulence and morphed into an unexpected, persistent, high dihedral configuration. As a result of the persistent high dihedral, the aircraft became unstable in a very divergent pitch mode in which the airspeed excursions from the nominal flight speed about doubled every cycle of the oscillation. The aircraft s design airspeed was subsequently exceeded and the resulting high dynamic pressures caused the wing leading edge secondary structure on the outer wing panels to fail and the solar cells and skin on the upper surface of the wing to rip away. As a result, the vehicle lost its ability to maintain lift, fell into the Pacific Ocean within the confines of the U.S. Navy's Pacific Missile Range Facility, and was destroyed. This paper describes the mishap and its causes, and presents the technical recommendations and lessons learned for improving the design
Jacob, Jamey; Mitchell, Taylor; Whyte, Seabrook
To facilitate safe storage of greenhouse gases such as CO2 and CH4, airborne monitoring is investigated. Conventional soil gas monitoring has difficulty in distinguishing gas flux signals from leakage with those associated with meteorologically driven changes. A low-cost, lightweight sensor system has been developed and implemented onboard a small unmanned aircraft that measures gas concentration and is combined with other atmospheric diagnostics, including thermodynamic data and velocity from hot-wire and multi-hole probes. To characterize the system behavior and verify its effectiveness, field tests have been conducted over controlled rangeland burns and over simulated leaks. In the former case, since fire produces carbon dioxide over a large area, this was an opportunity to test in an environment that while only vaguely similar to a carbon sequestration leak source, also exhibits interesting plume behavior. In the simulated field tests, compressed gas tanks are used to mimic leaks and generate gaseous plumes. Since the sensor response time is a function of vehicle airspeed, dynamic calibration models are required to determine accurate location of gas concentration in (x , y , z , t) . Results are compared with simulations using combined flight and atmospheric dynamic models. Supported by Department of Energy Award DE-FE0012173.
Highlights: • Heat accumulation in PCM causes failures of passive thermal management systems. • The introduction of forced air convection improves the reliability of PCMs. • Temperature distribution in the hybrid system remains uniform. • Active cooling and PCMs play separate roles in battery thermal management. • Numerical results agree with experiment data and give theoretic insights. - Abstract: Passive thermal management systems using phase change materials (PCMs) provides an effective solution to the overheating of lithium ion batteries. But this study shows heat accumulation in PCMs caused by the inefficient cooling of air natural convection leads to thermal management system failures: The temperature in a battery pack operating continuously outranges the safety limit of 60 °C after two cycles with discharge rate of 1.5 C and 2 C. Here a hybrid system that integrates PCMs with forced air convection is presented. This combined system successfully prevents heat accumulation and maintains the maximum temperature under 50 °C in all cycles. Study on airspeed effects reveals that thermo-physical properties of PCMs dictate the maximum temperature rise and temperature uniformity in the battery pack, while forced air convection plays a critical role in recovering thermal energy storage capacity of PCMs. A numerical study is also carried out and validated with experiment data, which gives theoretic insights on thermo-physical changes in this hybrid battery thermal management system
Halloran, Siobhan; Wexler, Anthony; Ristenpart, William
Virologists and other researchers who test pathogens for airborne disease transmissibility often place a test animal downstream from an inoculated animal and later determine whether the test animal became infected. Despite the crucial role of the airflow in modulating the pathogen transmission, to date the infectious disease community has paid little attention to the effect of airspeed or turbulence intensity on the probability of transmission. Here we present measurements of the turbulent dispersivity under conditions relevant to experimental tests of airborne disease transmissibility between laboratory animals. We used time lapse photography to visualize the downstream transport and turbulent dispersion of smoke particulates released from a point source downstream of a standard axial fan, thus mimicking the release and transport of expiratory aerosols exhaled by an inoculated animal. We demonstrate that the fan speed counterintuitively has no effect on the downstream plume width, a result replicated with a variety of different fan types and configurations. The results point toward a useful simplification in modeling of airborne disease transmission via fan-generated flows.
Human coughing is studied non-intrusively by high-speed schlieren videography, revealing a turbulent jet lasting up to 1 sec with a total expelled air volume of about 2 L. Velocimetry of eddy motion reveals a jet centerline airspeed of at least 8 m/sec. With Re roughly 18,000 the cough jet is inertia-driven and buoyancy is negligible. It shows typical round-turbulent-jet behavior, including a conical spreading angle of 24 deg, despite irregular initial conditions. The cough jet is projected several m into the surrounding air before it mixes out. It is well known that a cough can transmit infectious agents, and we are advised to cover our mouths in an apparent attempt to thwart the jet formation. Present experiments have shown that wearing a surgical mask or respirator designed to prevent the inhalation of infectious agents also interferes with the cough-jet formation, redirecting it into the person's rising thermal plume. (Tang et al., J. Royal. Soc. Interface 6, S727, 2009.)
Watts, Michael E.; Greenwood, Eric; Stephenson, James
This paper presents an overview of a flight test campaign performed at different test sites whose altitudes ranged from 0 to 7000 feet above mean sea level (AMSL) between September 2014 and February 2015. The purposes of this campaign were to: investigate the effects of altitude variation on noise generation, investigate the effects of gross weight variation on noise generation, establish the statistical variability in acoustic flight testing of helicopters, and characterize the effects of transient maneuvers on radiated noise for a medium-lift utility helicopter. In addition to describing the test campaign, results of the acoustic effects of altitude variation for the AS350 SD1 and EH-60L aircraft are presented. Large changes in acoustic amplitudes were observed in response to changes in ambient conditions when the helicopter was flown at constant indicated airspeed and gross weight at the three test sites. However, acoustic amplitudes were found to scale with ambient pressure when flight conditions were defined in terms of the non-dimensional parameters, such as the weight coefficient and effective hover tip Mach number.
Sim, Ben W.; Janakiram, Ram D.; Barbely, Natasha L.; Solis, Eduardo
Results from a recent joint DARPA/Boeing/NASA/Army wind tunnel test demonstrated the ability to reduce in-plane, low frequency noise of the full-scale Boeing-SMART rotor using active flaps. Test data reported in this paper illustrated that acoustic energy in the first six blade-passing harmonics could be reduced by up to 6 decibels at a moderate airspeed, level flight condition corresponding to advance ratio of 0.30. Reduced noise levels were attributed to selective active flap schedules that modified in-plane blade airloads on the advancing side of the rotor, in a manner, which generated counteracting acoustic pulses that partially offset the negative pressure peaks associated with in-plane, steady thickness noise. These favorable reduced-noise operating states are a strong function of the active flap actuation amplitude, frequency and phase. The associated noise reductions resulted in reduced aural detection distance by up to 18%, but incurred significant vibratory load penalties due to increased hub shear forces. Small reductions in rotor lift-to-drag ratios, of no more than 3%, were also measured
Kopsaftopoulos, Fotios; Nardari, Raphael; Li, Yu-Hung; Wang, Pengchuan; Chang, Fu-Kuo
In this work, the system design, integration, and wind tunnel experimental evaluation are presented for a bioinspired self-sensing intelligent composite unmanned aerial vehicle (UAV) wing. A total of 148 micro-sensors, including piezoelectric, strain, and temperature sensors, in the form of stretchable sensor networks are embedded in the layup of a composite wing in order to enable its self-sensing capabilities. Novel stochastic system identification techniques based on time series models and statistical parameter estimation are employed in order to accurately interpret the sensing data and extract real-time information on the coupled air flow-structural dynamics. Special emphasis is given to the wind tunnel experimental assessment under various flight conditions defined by multiple airspeeds and angles of attack. A novel modeling approach based on the recently introduced Vector-dependent Functionally Pooled (VFP) model structure is employed for the stochastic identification of the "global" coupled airflow-structural dynamics of the wing and their correlation with dynamic utter and stall. The obtained results demonstrate the successful system-level integration and effectiveness of the stochastic identification approach, thus opening new perspectives for the state sensing and awareness capabilities of the next generation of "fly-by-fee" UAVs.
McCabe, Jennifer D; Olsen, Brian J; Hiebeler, David
Suture zones are areas where range contact zones and hybrid zones of multiple taxa are clustered. Migratory divides, contact zones between divergent populations that breed adjacent to one another but use different migratory routes, are a particular case of suture zones. Although multiple hypotheses for both the formation and maintenance of migratory divides have been suggested, quantitative tests are scarce. Here, we tested whether a novel factor, prevailing winds, was sufficient to explain both the evolution and maintenance of the Cordilleran migratory divide using individual-based models. Empirical observations of eastern birds suggest a circuitous migratory route across Canada before heading south. Western breeders, however, travel south along the Pacific coast to their wintering grounds. We modeled the effect of wind on bird migratory flights by allowing them to float at elevation using spatially explicit modeled wind data. Modeled eastern birds had easterly mean trajectories, whereas western breeders showed significantly more southern trajectories. We also determined that a mean airspeed of 18.5 m s(-1) would be necessary to eliminate this difference in trajectory, a speed that is achieved by waterfowl and shorebirds, but is faster than songbird flight speeds. These results lend support for the potential importance of wind in shaping the phylogeographic history of North American songbirds.
Norzailawati, M. N.; Alias, A.; Akma, R. S.
This paper discusses on-going research related to zoning regulation for the remote sensing drone in the urban applications. Timestamped maps are presented here follow a citation-based approach, where significant information is retrieved from the scientific literature. The emergence of drones in domestic air raises lots understandable issues on privacy, security and uncontrolled pervasive surveillance that require a careful and alternative solution. The effective solution is to adopt a privacy and property rights approach that create a drone zoning and clear drone legislatures. In providing a differential trend to other reviews, this paper is not limited to drones zoning and regulations, but also, discuss on trend remote sensing drones specification in designing a drone zones. Remote sensing drone will specific according to their features and performances; size and endurance, maximum airspeed and altitude level and particular references are made to the drones range. The implementation of laws zoning could lie with the urban planners whereby, a zoning for drone could become a new tactic used to specify areas, where drones could be used, will provide remedies for the harm that arise from drones, and act as a different against irresponsible behaviour. Finally, underlines the need for next regulations on guidelines and standards which can be used as a guidance for urban decision makers to control the drones' operating, thus ensuring a quality and sustainability of resilience cities simultaneously encouraging the revolution of technology.
Liu Zhi; Wang Yong
Motivated by the autopilot of an unmanned aerial vehicle (UAV) with a wide flight enve-lope span experiencing large parametric variations in the presence of uncertainties, a fuzzy adaptive tracking controller (FATC) is proposed. The controller consists of a fuzzy baseline controller and an adaptive increment, and the main highlight is that the fuzzy baseline controller and adaptation laws are both based on the fuzzy multiple Lyapunov function approach, which helps to reduce the conservatism for the large envelope and guarantees satisfactory tracking performances with strong robustness simultaneously within the whole envelope. The constraint condition of the fuzzy baseline controller is provided in the form of linear matrix inequality (LMI), and it specifies the satisfactory tracking performances in the absence of uncertainties. The adaptive increment ensures the uniformly ultimately bounded (UUB) predication errors to recover satisfactory responses in the presence of uncertainties. Simulation results show that the proposed controller helps to achieve high-accuracy tracking of airspeed and altitude desirable commands with strong robustness to uncertainties throughout the entire flight envelope.
Hilton, D. A.; Pegg, R. J.
Noise measurements under controlled conditions have been made inside and outside of a school building during flyover operations of four different helicopters. The helicopters were operated at a condition considered typical for a police patrol mission. Flyovers were made at an altitude of 500 ft and an airspeed of 45 miles per hour. During these operations acoustic measurements were made inside and outside of the school building with the windows closed and then open. The outside noise measurements during helicopter flyovers indicate that the outside db(A) levels were approximately the same for all test helicopters. For the windows closed case, significant reductions for the inside measured db(A) values were noted for all overflights. These reductions were approximately 20 db(A); similar reductions were noted in other subjective measuring units. The measured internal db(A) levels with the windows open exceeded published classroom noise criteria values; however, for the windows-closed case they are in general agreement with the criteria values.
王如文; 郑来昌; 杨小辉; 杨克
制备了一种负载型Re基催化剂Re2O7/γ-Al2O3,用于直链内烯烃与乙烯歧化制备α-烯烃.结果表明,以C11～C12直链内烯烃为原料,反应温度60 ℃,反应体积空速1 h-1,反应压力3 MPa的条件下,C11～C12烯烃的单程转化率达到90.0％,歧化选择性达到85.98％.%Preparation of supported Re based catalysts modified Re2O7/λ-Al2O3 ,used in the preparation of a-olefin metathesis of olefins with ethylene in the straight chain internal olefins process. The results showed that to C11,to C12,a straight chain internal olefin as raw material, the reaction temperature of 60 ℃ ,the reaction volume airspeed 1 h-1, C11～C12 ,olefin-way conversion rate of 90.0% under the conditions of the reaction pressure 3 MPa,metathesis selectivity 85.98%.
Wright, Stephen; O'Hare, David
The analog dials in traditional GA aircraft cockpits are being replaced by integrated electronic displays, commonly referred to as glass cockpits. Pilots may be trained on glass cockpit aircraft or encounter them after training on traditional displays. The effects of glass cockpit displays on initial performance and potential transfer effects between cockpit display configurations have yet to be adequately investigated. Flight-naïve participants were trained on either a simulated traditional display cockpit or a simulated glass display cockpit. Flight performance was measured in a test flight using either the same or different cockpit display. Loss of control events and accuracy in controlling altitude, airspeed and heading, workload, and situational awareness were assessed. Preferences for cockpit display configurations and opinions on ease of use were also measured. The results revealed consistently poorer performance on the test flight for participants using the glass cockpit compared to the traditional cockpit. In contrast the post-flight questionnaire data revealed a strong subjective preference for the glass cockpit over the traditional cockpit displays. There was only a weak effect of prior training. The specific glass cockpit display used in this study was subjectively appealing but yielded poorer flight performance in participants with no previous flight experience than a traditional display. Performance data can contradict opinion data. The design of glass cockpit displays may present some difficulties for pilots in the very early stages of training.
Yu, Yan S. W.; Graff, Matthew M.; Bresee, Chris S.; Man, Yan B.; Hartmann, Mitra J. Z.
Observation of terrestrial mammals suggests that they can follow the wind (anemotaxis), but the sensory cues underlying this ability have not been studied. We identify a significant contribution to anemotaxis mediated by whiskers (vibrissae), a modality previously studied only in the context of direct tactile contact. Five rats trained on a five-alternative forced-choice airflow localization task exhibited significant performance decrements after vibrissal removal. In contrast, vibrissal removal did not disrupt the performance of control animals trained to localize a light source. The performance decrement of individual rats was related to their airspeed threshold for successful localization: animals that found the task more challenging relied more on the vibrissae for localization cues. Following vibrissal removal, the rats deviated more from the straight-line path to the air source, choosing sources farther from the correct location. Our results indicate that rats can perform anemotaxis and that whiskers greatly facilitate this ability. Because air currents carry information about both odor content and location, these findings are discussed in terms of the adaptive significance of the interaction between sniffing and whisking in rodents. PMID:27574705
It has been hypothesized that a human pilot uses the same set of generic skills to control a wide variety of aircraft. If this is true, then it should be possible to construct an electronic controller which embodies this generic skill set such that it can successfully control different airplanes without being matched to a specific airplane. In an attempt to create such a system, a fuzzy logic controller was devised to control aileron or roll spoiler position. This controller was used to control bank angle for both a piston powered single engine aileron equipped airplane simulation and a business jet simulation which used spoilers for primary roll control. Overspeed, stall and overbank protection were incorporated in the form of expert systems supervisors and weighted fuzzy rules. It was found that by using the artificial intelligence techniques of fuzzy logic and expert systems, a generic lateral controller could be successfully used on two general aviation aircraft types that have very different characteristics. These controllers worked for both airplanes over their entire flight envelopes. The controllers for both airplanes were identical except for airplane specific limits (maximum allowable airspeed, throttle ]ever travel, etc.). This research validated the fact that the same fuzzy logic based controller can control two very different general aviation airplanes. It also developed the basic controller architecture and specific control parameters required for such a general controller.
Bull, Daniel Mark
The purpose of this thesis was to conduct preliminary research, in the form of a pilot study, concerning the natural effects of hypoxia compared to the effects of hypoxia experienced after the consumption of an energy beverage. The study evaluated the effects of hypoxia on FAA certificated pilots at a simulated legal general aviation altitude, utilizing the normobaric High Altitude Lab (HAL) located at Embry Riddle Aeronautical University, Daytona Beach, Florida. The researcher tested 11 subjects, who completed three simulated flight tasks within the HAL using the Frasca International Mentor Advanced Aviation Training Device (AATD). The flight tasks were completed after consuming Red BullRTM, MonsterRTM , or a placebo beverage. The researcher derived three test variables from core outputs of the AATD: lateral deviations from the glide slope, vertical deviations from the localizer, and airspeed deviations from the target speed of 100 knots. A repeated-measures ANOVA was carried out to determine effects of the beverages on the test variables. While results were non-significant, the researcher concluded that further research should be conducted with a larger sample.
Belcastro, C. M.; Ostroff, A. J.
Low-altitude wind shear is recognized as an infrequent but significant hazard to all aircraft during take-off and landing. A total energy-rate sensor, which is potentially applicable to this problem, has been developed for measuring specific total energy-rate of an airplane with respect to the air mass. This paper presents control system designs, with and without energy-rate feedback, for the approach to landing of a transport airplane through severe wind shear and gusts to evaluate application of this sensor. A system model is developed which incorporates wind shear dynamics equations with the airplance equations of motion, thus allowing the control systems to be analyzed under various wind shears. The control systems are designed using optimal output feedback and are analyzed using frequency domain control theory techniques. Control system performance is evaluated using a complete nonlinear simulation of the airplane and a severe wind shear and gust data package. The analysis and simulation results indicate very similar stability and performance characteristics for the two designs. An implementation technique for distributing the velocity gains between airspeed and ground speed in the simulation is also presented, and this technique is shown to improve the performance characteristics of both designs.
A. R. Rodi
Full Text Available Geometric altitude data from a combined Global Navigation Satellite System (GNSS and inertial measurement unit (IMU system on the University of Wyoming King Air research aircraft are used to estimate acceleration effects on static pressure measurement. Using data collected during periods of accelerated flight, comparison of measured pressure with that derived from GNSS/IMU geometric altitude show that errors exceeding 150 Pa can occur which is significant in airspeed and atmospheric air motion determination. A method is developed to predict static pressure errors from analysis of differential pressure measurements from a Rosemount model 858 differential pressure air velocity probe. The method was evaluated with a carefully designed probe towed on connecting tubing behind the aircraft – a "trailing cone" – in steady flight, and shown to have a precision of about ±10 Pa over a wide range of conditions including various altitudes, power settings, and gear and flap extensions. Under accelerated flight conditions, compared to the GNSS/IMU data, this algorithm predicts corrections to a precision of better than ±20 Pa. Some limiting factors affecting the precision of static pressure measurement on a research aircraft are examined.
Greenwood, Eric; Sim, Ben W.; Boyd, D. Douglas, Jr.
The effects of ambient atmospheric conditions, air temperature and density, on rotor harmonic noise radiation are characterized using theoretical models and experimental measurements of helicopter noise collected at three different test sites at elevations ranging from sea level to 7000 ft above sea level. Significant changes in the thickness, loading, and blade-vortex interaction noise levels and radiation directions are observed across the different test sites for an AS350 helicopter flying at the same indicated airspeed and gross weight. However, the radiated noise is shown to scale with ambient pressure when the flight condition of the helicopter is defined in nondimensional terms. Although the effective tip Mach number is identified as the primary governing parameter for thickness noise, the nondimensional weight coefficient also impacts lower harmonic loading noise levels, which contribute strongly to low frequency harmonic noise radiation both in and out of the plane of the horizon. Strategies for maintaining the same nondimensional rotor operating condition under different ambient conditions are developed using an analytical model of single main rotor helicopter trim and confirmed using a CAMRAD II model of the AS350 helicopter. The ability of the Fundamental Rotorcraft Acoustics Modeling from Experiments (FRAME) technique to generalize noise measurements made under one set of ambient conditions to make accurate noise predictions under other ambient conditions is also validated.
Gilbert, Michael G.; Silva, Walter A.
A new design concept in the development of VTOL aircraft with high forward flight speed capability is that of the X-Wing, a stiff, bearingless helicopter rotor system which can be stopped in flight and the blades used as two forward-swept and two aft-swept wings. Because of the usual configuration in the fixed-wing mode, there is a high potential for aeroelastic divergence or flutter and coupling of blade vibration modes with rigid-body modes. An aeroelastic stability analysis of an X-Wing configuration aircraft was undertaken to determine if these problems could exist. This paper reports on the results of dynamic stability analyses in the lateral and longitudinal directions including the vehicle rigid-body and flexible modes. A static aeroelastic analysis using the normal vibration mode equations of motion was performed to determine the cause of a loss of longitudinal static margin with increasing airspeed. This loss of static margin was found to be due to aeroelastic washin of the forward-swept blades and washout of the aft-swept blades moving the aircraft aerodynamic center forward of the center of gravity. This phenomenon is likely to be generic to X-Wing aircraft.
Martos, Borja; Kiszely, Paul; Foster, John V.
As part of the NASA Aviation Safety Program (AvSP), a novel pitot-static calibration method was developed to allow rapid in-flight calibration for subscale aircraft while flying within confined test areas. This approach uses Global Positioning System (GPS) technology coupled with modern system identification methods that rapidly computes optimal pressure error models over a range of airspeed with defined confidence bounds. This method has been demonstrated in subscale flight tests and has shown small 2- error bounds with significant reduction in test time compared to other methods. The current research was motivated by the desire to further evaluate and develop this method for full-scale aircraft. A goal of this research was to develop an accurate calibration method that enables reductions in test equipment and flight time, thus reducing costs. The approach involved analysis of data acquisition requirements, development of efficient flight patterns, and analysis of pressure error models based on system identification methods. Flight tests were conducted at The University of Tennessee Space Institute (UTSI) utilizing an instrumented Piper Navajo research aircraft. In addition, the UTSI engineering flight simulator was used to investigate test maneuver requirements and handling qualities issues associated with this technique. This paper provides a summary of piloted simulation and flight test results that illustrates the performance and capabilities of the NASA calibration method. Discussion of maneuver requirements and data analysis methods is included as well as recommendations for piloting technique.
Kottapalli, Sesi; Hagerty, Brandon; Salazar, Denise
A full-scale helicopter smart material actuated rotor technology (SMART) rotor test was conducted in the USAF National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames. The SMART rotor system is a five-bladed MD 902 bearingless rotor with active trailing-edge flaps. The flaps are actuated using piezoelectric actuators. Rotor performance, structural loads, and acoustic data were obtained over a wide range of rotor shaft angles of attack, thrust, and airspeeds. The primary test objective was to acquire unique validation data for the high-performance computing analyses developed under the Defense Advanced Research Project Agency (DARPA) Helicopter Quieting Program (HQP). Other research objectives included quantifying the ability of the on-blade flaps to achieve vibration reduction, rotor smoothing, and performance improvements. This data set of rotor performance and structural loads can be used for analytical and experimental comparison studies with other full-scale rotor systems and for analytical validation of computer simulation models. The purpose of this final data report is to document a comprehensive, highquality data set that includes only data points where the flap was actively controlled and each of the five flaps behaved in a similar manner.
Lobitz, D.W.; Veers, P.S.
As the technology for horizontal axis wind turbines (HAWT) development matures, more novel techniques are required for the capture of additional amounts of energy, alleviation of loads and control of the rotor. One such technique employs the use of an adaptive blade that could sense the wind velocity or rotational speed in some fashion and accordingly modify its aerodynamic configuration to meet a desired objective. This could be achieved in either an active or passive manner, although the passive approach is much more attractive due to its simplicity and economy. As an example, a blade design might employ coupling between bending and/or extension, and twisting so that, as it bends and extends due to the action of the aerodynamic and inertial loads, it also twists modifying the aerodynamic performance in some way. These performance modifications also have associated aeroelastic effects, including effects on aeroelastic instability. To address the scope and magnitude of these effects a tool has been developed for investigating classical flutter and divergence of HAWT blades. As a starting point, an adaptive version of the uniform Combined Experiment Blade will be investigated. Flutter and divergence airspeeds will be reported as a function of the strength of the coupling and also be compared to those of generic blade counterparts.
Stell, Laurel L.
To enable arriving aircraft to fly optimized descents computed by the flight management system (FMS) in congested airspace, ground automation must accurately predict descent trajectories. To support development of the trajectory predictor and its error models, commercial flights executed idle-thrust descents, and the recorded data includes the target speed profile and FMS intent trajectories. The FMS computes the intended descent path assuming idle thrust after top of descent (TOD), and any intervention by the controllers that alters the FMS execution of the descent is recorded so that such flights are discarded from the analysis. The horizontal flight path, cruise and meter fix altitudes, and actual TOD location are extracted from the radar data. Using more than 60 descents in Boeing 777 aircraft, the actual speeds are compared to the intended descent speed profile. In addition, three aspects of the accuracy of the FMS intent trajectory are analyzed: the meter fix crossing time, the TOD location, and the altitude at the meter fix. The actual TOD location is within 5 nmi of the intent location for over 95% of the descents. Roughly 90% of the time, the airspeed is within 0.01 of the target Mach number and within 10 KCAS of the target descent CAS, but the meter fix crossing time is only within 50 sec of the time computed by the FMS. Overall, the aircraft seem to be executing the descents as intended by the designers of the onboard automation.
Ali, Syed Firasat; Khan, Javed Khan; Rossi, Marcia J.; Crane, Peter; Heath, Bruce E.; Knighten, Tremaine; Culpepper, Christi
Personal computer based flight simulators are expanding opportunities for providing low-cost pilot training. One advantage of these devices is the opportunity to incorporate instructional features into training scenarios that might not be cost effective with earlier systems. Research was conducted to evaluate the utility of different instructional features using a coordinated level turn as an aircraft maneuvering task. In study I, a comparison was made between automated computer grades of performance with certified flight instructors grades. Every one of the six student volunteers conducted a flight with level turns at two different bank angles. The automated computer grades were based on prescribed tolerances on bank angle, airspeed and altitude. Two certified flight instructors independently examined the video tapes of heads up and instrument displays of the flights and graded them. The comparison of automated grades with the instructors grades was based on correlations between them. In study II, a 2x2 between subjects factorial design was used to devise and conduct an experiment. Comparison was made between real time training and above real time training and between feedback and no feedback in training. The performance measure to monitor progress in training was based on deviations in bank angle and altitude. The performance measure was developed after completion of the experiment including the training and test flights. It was not envisaged before the experiment. The experiment did not include self- instructions as it was originally planned, although feedback by experimenter to the trainee was included in the study.
J. P. Fugal
Full Text Available Holographic data from the prototype airborne digital holographic instrument HOLODEC (Holographic Detector for Clouds, taken during test flights are digitally reconstructed to obtain the size (equivalent diameters in the range 23 to 1000 μm, three-dimensional position, and two-dimensional profile of ice particles and then ice particle size distributions and number densities are calculated using an automated algorithm with minimal user intervention. The holographic method offers the advantages of a well-defined sample volume size that is not dependent on particle size or airspeed, and offers a unique method of detecting shattered particles. The holographic method also allows the volume sample rate to be increased beyond that of the prototype HOLODEC instrument, limited solely by camera technology.
HOLODEC size distributions taken in mixed-phase regions of cloud compare well to size distributions from a PMS FSSP probe also onboard the aircraft during the test flights. A conservative algorithm for detecting shattered particles utilizing the particles depth-position along the optical axis eliminates the obvious ice particle shattering events from the data set. In this particular case, the size distributions of non-shattered particles are reduced by approximately a factor of two for particles 15 to 70 μm in equivalent diameter, compared to size distributions of all particles.
Kottapalli, Sesi B. R.
Measured, open loop and closed loop data from the SMART rotor test in the NASA Ames 40- by 80- Foot Wind Tunnel are compared with CAMRAD II calculations. One open loop high-speed case and four closed loop cases are considered. The closed loop cases include three high-speed cases and one low-speed case. Two of these high-speed cases include a 2 deg flap deflection at 5P case and a test maximum-airspeed case. This study follows a recent, open loop correlation effort that used a simple correction factor for the airfoil pitching moment Mach number. Compared to the earlier effort, the current open loop study considers more fundamental corrections based on advancing blade aerodynamic conditions. The airfoil tables themselves have been studied. Selected modifications to the HH-06 section flap airfoil pitching moment table are implemented. For the closed loop condition, the effect of the flap actuator is modeled by increased flap hinge stiffness. Overall, the open loop correlation is reasonable, thus confirming the basic correctness of the current semi-empirical modifications; the closed loop correlation is also reasonable considering that the current flap model is a first generation model. Detailed correlation results are given in the paper.
Ustinov, Eugene A.
A concept of an aero-assisted pre-stage is proposed, which enables launch of both ballistic and aero-assisted launch vehicles from conventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle, so that their flight directions are coaligned. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensure prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet airliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achieved, and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is activated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aero-assisted vehicles (spaceplanes) are considered.
Mccarthy, J.; Blick, E. F.; Bensch, R. R.
Several hours of three dimensional wind data were collected in the thunderstorm approach-to-landing environment, using an instrumented Queen Air airplane. These data were used as input to a numerical simulation of aircraft response, concentrating on fixed-stick assumptions, while the aircraft simulated an instrument landing systems approach. Output included airspeed, vertical displacement, pitch angle, and a special approach deterioration parameter. Theory and the results of approximately 1000 simulations indicated that about 20 percent of the cases contained serious wind shear conditions capable of causing a critical deterioration of the approach. In particular, the presence of high energy at the airplane's phugoid frequency was found to have a deleterious effect on approach quality. Oscillations of the horizontal wind at the phugoid frequency were found to have a more serious effect than vertical wind. A simulation of Eastern flight 66, which crashed at JFK in 1975, served to illustrate the points of the research. A concept of a real-time wind shear detector was outlined utilizing these results.
This paper presents a novel bat-like unmanned aerial vehicle inspired by the morphing-wing mechanism of bats. The goal of this paper is twofold. Firstly, a modelling framework is introduced for analysing how the robot should manoeuvre by means of changing wing morphology. This allows the definition of requirements for achieving forward and turning flight according to the kinematics of the wing modulation. Secondly, an attitude controller named backstepping+DAF is proposed. Motivated by biological evidence about the influence of wing inertia on the production of body accelerations, the attitude control law incorporates wing inertia information to produce desired roll (φ) and pitch (θ) acceleration commands (desired angular acceleration function (DAF)). This novel control approach is aimed at incrementing net body forces (Fnet) that generate propulsion. Simulations and wind-tunnel experimental results have shown an increase of about 23% in net body force production during the wingbeat cycle when the wings are modulated using the DAF as a part of the backstepping control law. Results also confirm accurate attitude tracking in spite of high external disturbances generated by aerodynamic loads at airspeeds up to 5 ms−1. (paper)
Hochstetler, R.D. [Research Adventures,Inc., Kensington, MD (United States)
In recent years the study of Earth processes has increased significantly. Conventional aircraft have been employed to a large extent in gathering much of this information. However, with this expansion of research has come the need to investigate and measure phenomena that occur beyond the performance capabilities of conventional aircraft. Where long dwell times or observations at very low attitudes are required there are few platforms that can operate safely, efficiently, and cost-effectively. One type of aircraft that meets all three parameters is the unmanned, autonomously operated airship. The UAV airship is smaller than manned airships but has similar performance characteristics. It`s low speed stability permits high resolution observations and provides a low vibration environment for motion sensitive instruments. Maximum airspeed is usually 30mph to 35mph and endurance can be as high as 36 hours. With scientific payload capacities of 100 kilos and more, the UAV airship offers a unique opportunity for carrying significant instrument loads for protracted periods at the air/surface interface. The US Army has operated UAV airships for several years conducting border surveillance and monitoring, environmental surveys, and detection and mapping of unexploded ordinance. The technical details of UAV airships, their performance, and the potential of such platforms for more advanced research roles will be presented. 3 refs., 5 figs.
Full Text Available This paper presents a complete procedure for sensor compatibility correction of a fixed-wing Unmanned Air Vehicle (UAV. The sensors consist of a differential air pressure transducer for airspeed measurement, two airdata vanes installed on an airdata probe for angle of attack (AoA and angle of sideslip (AoS measurement, and an Attitude and Heading Reference System (AHRS that provides attitude angles, angular rates, and acceleration. The procedure is mainly based on a two pass algorithm called the Rauch-Tung-Striebel (RTS smoother, which consists of a forward pass Extended Kalman Filter (EKF and a backward recursion smoother. On top of that, this paper proposes the implementation of the Wiener Type Filter prior to the RTS in order to avoid the complicated process noise covariance matrix estimation. Furthermore, an easy to implement airdata measurement noise variance estimation method is introduced. The method estimates the airdata and subsequently the noise variances using the ground speed and ascent rate provided by the Global Positioning System (GPS. It incorporates the idea of data regionality by assuming that some sort of statistical relation exists between nearby data points. Root mean square deviation (RMSD is being employed to justify the sensor compatibility. The result shows that the presented procedure is easy to implement and it improves the UAV sensor data compatibility significantly.
Chan, Woei-Leong; Hsiao, Fei-Bin
This paper presents a complete procedure for sensor compatibility correction of a fixed-wing Unmanned Air Vehicle (UAV). The sensors consist of a differential air pressure transducer for airspeed measurement, two airdata vanes installed on an airdata probe for angle of attack (AoA) and angle of sideslip (AoS) measurement, and an Attitude and Heading Reference System (AHRS) that provides attitude angles, angular rates, and acceleration. The procedure is mainly based on a two pass algorithm called the Rauch-Tung-Striebel (RTS) smoother, which consists of a forward pass Extended Kalman Filter (EKF) and a backward recursion smoother. On top of that, this paper proposes the implementation of the Wiener Type Filter prior to the RTS in order to avoid the complicated process noise covariance matrix estimation. Furthermore, an easy to implement airdata measurement noise variance estimation method is introduced. The method estimates the airdata and subsequently the noise variances using the ground speed and ascent rate provided by the Global Positioning System (GPS). It incorporates the idea of data regionality by assuming that some sort of statistical relation exists between nearby data points. Root mean square deviation (RMSD) is being employed to justify the sensor compatibility. The result shows that the presented procedure is easy to implement and it improves the UAV sensor data compatibility significantly. PMID:22163819
Reynolds, Kate V.; Thomas, Adrian L. R.; Taylor, Graham K.
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. PMID:25320064
Wang, Wei; Li, Aijun; Xie, Yanwu; Tan, Jian
Historically, aircraft longitudinal control has been realized by means of two loops: flight path (the control variable is elevator displacement) and speed control (the control variable is propulsive thrust or engine power). Both the elevator and throttle control cause coupled altitude and speed response, which exerts negative effects on longitudinal flight performance of aircraft, especially for Terrain Following(TF) flight. Energy-based method can resolve coupled problem between flight speed and path by controlling total energy rate and energy distribution rate between elevator and throttle. In this paper, energy-based control method is applied to design a TF flight control system for controlling flight altitude directly. An error control method of airspeed and altitude is adopted to eliminate the stable error of the total energy control system when decoupling control. Pitch loop and pitch rate feedback loop are designed for the system to damp the oscillatory response produced by TF system. The TF flight control system structure diagram and an aircraft point-mass energy motion model including basic control loops are given and used to simulate decoupling performance of the TF fight control system. Simulation results show that the energy-based TF flight control system can decouple flight velocity and flight path angle, exactly follow planned flight path, and greatly reduce altitude error, which is between +10m and -8m.
Hueschen, Richard M.; Khong, Thuan H.
A vertical navigation (VNAV) outer-loop control system was developed to capture and track the vertical path segments of energy-efficient trajectories that are being developed for high-density operations in the evolving Next Generation Air Transportation System (NextGen). The VNAV control system has a speed-on-elevator control mode to pitch the aircraft for tracking a calibrated airspeed (CAS) or Mach number profile and a path control mode for tracking the VNAV altitude profile. Mode control logic was developed for engagement of either the speed or path control modes. The control system will level the aircraft to prevent it from flying through a constraint altitude. A stability analysis was performed that showed that the gain and phase margins of the VNAV control system significantly exceeded the design gain and phase margins. The system performance was assessed using a six-deg-of-freedom non-linear transport aircraft simulation and the performance is illustrated with time-history plots of recorded simulation data.
Full Text Available In this paper, we show that in mixed phase clouds FSSP-100 measurements may be contaminated by ice crystals, inducing wrong interpretation of particle size and subsequent bulk parameters. This contamination is generally revealed by a bimodal feature of the particle size distribution; in other words, in mixed phase clouds bimodal features could be an indication of the presence of ice particles. The combined measurements of the FSSP-100 and the Polar Nephelometer give a coherent description of the effect of the ice crystals on the FSSP-100 response. The FSSP-100 particle size distributions are characterized by a bimodal shape with a second mode peaked between 25 and 35 μm related to ice crystals. This feature is observed with the FSSP-100 at airspeed up to 200 m s−1 and with the FSSP-300 series. In order to assess the size calibration for clouds of ice crystals the response of the FSSP-100 probe has been numerically simulated using a light scattering model of randomly oriented hexagonal ice particles and assuming both smooth and rough crystal surfaces. The results suggest that the second mode measured between 25 μm and 35 μm, does not necessarily represent true size responses but likely corresponds to bigger aspherical ice particles. According to simulation results, the sizing understatement would be neglected in the rough case but would be major with the smooth case. Qualitatively, the Polar Nephelometer phase function suggests that the rough case is the more suitable to describe real crystals. Quantitatively, however, it is difficult to conclude. Previous cloud in situ measurements suggest that the FSSP-100 secondary mode, peaked in the range 25–35 μm, is likely to be due to the shattering of large ice crystals on the probe tips. This finding is supported by the rather good relationship between the concentration of particles larger than 20 μm (hypothesized to be ice shattered-fragments measured by the FSSP and the
Canacci, Victor A.; Gonsalez, Jose C.; Spera, David A.; Burke, Thomas (Technical Monitor)
Major modifications were made in 1999 to the 6- by 9-Foot (1.8- by 2.7-m) Icing Research tunnel (IRT) at the NASA Glenn Research Center, including replacement of its heat exchanger and associated ducts and turning vanes, and the addition of fan outlet guide vanes (OGV's). A one-tenth scale model of the IRT (designated as the SMIRT) was constructed with and without these modifications and tested to increase confidence in obtaining expected improvements in flow quality around the tunnel loop. The SMIRT is itself an aerodynamic test facility whose flow patterns without modifications have been shown to be accurate, scaled representations of those measured in the IRT prior to the 1999 upgrade program. In addition, tests in the SMIRT equipped with simulated OGV's indicated that these devices in the IRT might reduce flow distortions immediately downstream of the fan by two thirds. Flow quality parameters measured in the SMIRT were projected to the full-size modified IRT, and quantitative estimates of improvements in flow quality were given prior to construction. In this paper, the results of extensive flow quality studies conducted in the SMIRT are documented. Samples of these are then compared with equivalent measurements made in the full-scale IRT, both before and after its configuration was upgraded. Airspeed, turbulence intensity, and flow angularity distributions are presented for cross sections downstream of the drive fan, both upstream and downstream of the replacement flat heat exchanger, in the stilling chamber, in the test section, and in the wakes of the new comer turning vanes with their unique expanding and contracting designs. Lessons learned from these scale-model studies are discussed.
Gordon, Craig A.
This thesis examines the ability of a small, single-engine airplane to return to the runway following an engine failure shortly after takeoff. Two sets of trajectories are examined. One set of trajectories has the airplane fly a straight climb on the runway heading until engine failure. The other set of trajectories has the airplane perform a 90° turn at an altitude of 500 feet and continue until engine failure. Various combinations of wind speed, wind direction, and engine failure times are examined. The runway length required to complete the entire flight from the beginning of the takeoff roll to wheels stop following the return to the runway after engine failure is calculated for each case. The optimal trajectories following engine failure consist of three distinct segments: a turn back toward the runway using a large bank angle and angle of attack; a straight glide; and a reversal turn to align the airplane with the runway. The 90° turn results in much shorter required runway lengths at lower headwind speeds. At higher headwind speeds, both sets of trajectories are limited by the length of runway required for the landing rollout, but the straight climb cases generally require a lower angle of attack to complete the flight. The glide back to the runway is performed at an airspeed below the best glide speed of the airplane due to the need to conserve potential energy after the completion of the turn back toward the runway. The results are highly dependent on the rate of climb of the airplane during powered flight. The results of this study can aid the pilot in determining whether or not a return to the runway could be performed in the event of an engine failure given the specific wind conditions and runway length at the time of takeoff. The results can also guide the pilot in determining the takeoff profile that would offer the greatest advantage in returning to the runway.
杨跃能; 郑伟; 吴杰
研究了变化风场中无人机的动力学建模与飞行特性问题.给出了大气扰动的Von Kaman模型,推导出变化风场中无人机的六自由度非线性动力学模型,最后基于上述模型进行数值计算,分析了变化风场中无人机的飞行特性.结果表明,该模型能准确地描述无人机的动力学和运动学规律;大气扰动下,各运动参数均产生不同幅度的振荡;同一频谱的扰动在不同方向上对无人机的影响差异较大;地速航向角和空速航向角存在偏差,侧向扰动对航向的影响最为显著.%A nonlinear six degrees of freedom dynamics model of an unmanned aerial vehicle(UAV) with wind effects and its flight simulation analysis are presented. Firstly, the Von Kaman model of atmosphere disturbance is described. Then, the nonlinear six degrees of freedom dynamics model of the UAV is derived. Finally, flight simulations of the UAV under atmosphere disturbance are done. The results show that the model is valid and accurate.The atmosphere disturbance will ioduce vibration on motion parameters. Under the same power spectral density, the effects of atmosphere disturbance are different according to the orientations. Groundspeed heading is not identical with airspeed heading under atmosphere disturbance, and the effect of cross disturbance on heading is the most remarkable.
Full Text Available This study investigates whether the 3-D wind vector can be measured reliably from a highly transportable and low-cost weight-shift microlight aircraft. We draw up a transferable procedure to accommodate flow distortion originating from the aircraft body and -wing. This procedure consists of the analysis of aircraft dynamics and seven successive calibration steps. For our aircraft the horizontal wind components receive their greatest single amendment (14 %, relative to the initial uncertainty from the correction of flow distortion magnitude in the dynamic pressure computation. Conversely the vertical wind component is most of all improved (31 % by subsequent steps considering the 3-D flow distortion distribution in the flow angle computations. Therein the influences of the aircraft's trim (53 %, as well as changes in the aircraft lift (16 % are considered by using the measured lift coefficient as explanatory variable. Three independent lines of analysis are used to evaluate the quality of the wind measurement: (a A wind tunnel study in combination with the propagation of sensor uncertainties defines the systems input uncertainty to ≈0.6 m s−1 at the extremes of a 95 % confidence interval. (b During severe vertical flight manoeuvres the deviation range of the vertical wind component does not exceed 0.3 m s−1. (c The comparison with ground based wind measurements yields an overall operational uncertainty (root mean square error of ≈0.4 m s−1 for the horizontal and ≈0.3 m s−1 for the vertical wind components. No conclusive dependence of the uncertainty on the wind magnitude (<8 m s−1 or true airspeed (ranging from 23–30 m s−1 is found. Hence our analysis provides the necessary basis to study the wind measurement precision and spectral quality, which is prerequisite for reliable Eddy-Covariance flux measurements.
On a December night in 1995, 159 passengers and crewmembers died when American Airlines Flight 965 flew into the side of a mountain while in route to Cali, Colombia. A key factor in the tragedy: The pilots had lost situational awareness in the dark, unfamiliar terrain. They had no idea the plane was approaching a mountain until the ground proximity warning system sounded an alarm only seconds before impact. The accident was of the kind most common at the time CFIT, or controlled flight into terrain says Trey Arthur, research aerospace engineer in the Crew Systems and Aviation Operations Branch at NASA s Langley Research Center. In situations such as bad weather, fog, or nighttime flights, pilots would rely on airspeed, altitude, and other readings to get an accurate sense of location. Miscalculations and rapidly changing conditions could contribute to a fully functioning, in-control airplane flying into the ground. To improve aviation safety by enhancing pilots situational awareness even in poor visibility, NASA began exploring the possibilities of synthetic vision creating a graphical display of the outside terrain on a screen inside the cockpit. How do you display a mountain in the cockpit? You have to have a graphics-powered computer, a terrain database you can render, and an accurate navigation solution, says Arthur. In the mid-1990s, developing GPS technology offered a means for determining an aircraft s position in space with high accuracy, Arthur explains. As the necessary technologies to enable synthetic vision emerged, NASA turned to an industry partner to develop the terrain graphical engine and database for creating the virtual rendering of the outside environment.
谭健; 周洲; 祝小平; 许晓平
For the longitudinal landing control problem of flying⁃wing UAV with unknown external disturbances, a backstepping L2 gain robust control scheme based on super twisting sliding mode disturbance observer and tracking differentiator is proposed. The tracking differentiator is introduced to calculate the derivative of virtual control law which is very difficult to evaluate with the traditional backstepping control. Super twisting sliding mode disturbance observer and L2 gain robust item are designed to increase the robustness of the control system. Simulation results show:the altitude and airspeed of UAV are tracked on control command, vertical ground speed is within the allowable range. Compared with traditional PID control scheme, the proposed control scheme has better automatical landing control performance.%针对存在干扰的飞翼布局无人机纵向着陆控制问题，提出一种基于super twisting滑模干扰观测器与跟踪微分器的反步L2增益鲁棒控制方案。为解决反步控制虚拟控制量求导复杂的问题，设计了跟踪微分器对虚拟控制量进行求导，同时综合采用super twisting滑模干扰观测器和L2增益鲁棒项增强了控制系统的鲁棒性。仿真结果表明，无人机高度、空速都跟踪上控制指令，垂直接地速度在允许的范围内，与传统的PID着陆控制方案相比具有更好的着陆控制性能。
Robuck, Mark; Wilkerson, Joseph; Maciolek, Robert; Vonderwell, Dan
A multi-year study was conducted under NASA NNA06BC41C Task Order 10 and NASA NNA09DA56C task orders 2, 4, and 5 to identify the most promising propulsion system concepts that enable rotor cruise tip speeds down to 54% of the hover tip speed for a civil tiltrotor aircraft. Combinations of engine RPM reduction and 2-speed drive systems were evaluated. Three levels of engine and the drive system advanced technology were assessed; 2015, 2025 and 2035. Propulsion and drive system configurations that resulted in minimum vehicle gross weight were identified. Design variables included engine speed reduction, drive system speed reduction, technology, and rotor cruise propulsion efficiency. The NASA Large Civil Tiltrotor, LCTR, aircraft served as the base vehicle concept for this study and was resized for over thirty combinations of operating cruise RPM and technology level, quantifying LCTR2 Gross Weight, size, and mission fuel. Additional studies show design sensitivity to other mission ranges and design airspeeds, with corresponding relative estimated operational cost. The lightest vehicle gross weight solution consistently came from rotor cruise tip speeds between 422 fps and 500 fps. Nearly equivalent results were achieved with operating at reduced engine RPM with a single-speed drive system or with a two-speed drive system and 100% engine RPM. Projected performance for a 2025 engine technology provided improved fuel flow over a wide range of operating speeds relative to the 2015 technology, but increased engine weight nullified the improved fuel flow resulting in increased aircraft gross weights. The 2035 engine technology provided further fuel flow reduction and 25% lower engine weight, and the 2035 drive system technology provided a 12% reduction in drive system weight. In combination, the 2035 technologies reduced aircraft takeoff gross weight by 14% relative to the 2015 technologies.
Johanna Aristizábal Galvis
Full Text Available This study proposes and develops a new technique for dry-route dextrin production consisting of converting cassava starch pellets on a fixed-bed dryer; this technique is more applicable to rural Colombian agro-business in technical, economic, social and environmental terms, particularly to so-called “rallanderías” compared to currently available dextrin production technology. The proposed process is practically clean, requires low investment, allows humid starch-cake to be directly used without the need for expensive pre-drying equipment, eliminates large quantities of dust being produced thereby leading to an easily-handled and packaged product being obtained. Different dex-trinisation technologies were compared; a pilot-line was implemented which included blending, granulation and drying units. The variables evaluated were cassava-starch variety, catalyst concentration and agglutinant type and concentration; pellet-size, bed-thickness and air-speed were also evaluated during blending, granulation and drying stages, respectively. It was determined that using 0.1-0.3% HCl on cassava starch, 1.5-3% cassava starch paste, L/D=1.25 pellets, a 55ºC pre -drying phase and 150ºC final conversion on 2 cm thickness fixed-bed dryer at 2-3 m/s air speed led to obtaining low friability (13%, high rupture force (1.3 kg-f, high solubility (90-100% and low fluidity (50-70 s dextrin pellets. An adhesive was then obtained from the dextrin resulting from the process described above for sealing cardboard-boxes and cartons having greater stickiness, tensile strength and stability compared to corn dextrin adhesive, suggesting that the proposed new cassava dextrin production technique constitutes a good technological option for adding value to Colombian cassava production at small “rallandería” level.
Metzger, S.; Junkermann, W.; Butterbach-Bahl, K.; Schmid, H. P.; Foken, T.
This study investigates whether the 3-D wind vector can be measured reliably from a highly transportable and low-cost weight-shift microlight aircraft. Therefore we draw up a transferable procedure to accommodate flow distortion originating from the aircraft body and -wing. This procedure consists of the analysis of aircraft dynamics and seven successive calibration steps. For our aircraft the horizontal wind components receive their greatest single amendment (14 %, relative to the initial uncertainty) from the correction of flow distortion magnitude in the dynamic pressure computation. Conversely the vertical wind component is most of all improved (31 %) by subsequent steps considering the 3-D flow distortion distribution in the flow angle computations. Therein the influences of the aircraft's trim (53 %), as well as changes in the aircraft lift (16 %) are considered by using the measured lift coefficient as explanatory variable. Three independent lines of analysis are used to evaluate the quality of the wind measurement: (a) A wind tunnel study in combination with the propagation of sensor uncertainties defines the systems input uncertainty to ≈0.6 m s-1 at the extremes of a 95 % confidence interval. (b) During severe vertical flight manoeuvres the deviation range of the vertical wind component does not exceed 0.3 m s-1. (c) The comparison with ground based wind measurements yields an overall operational uncertainty (root mean square error) of ≈0.4 m s-1 for the horizontal and ≈0.3 m s-1 for the vertical wind components. No conclusive dependence of the uncertainty on the wind magnitude (<8 m s-1) or true airspeed (ranging from 23-30 m s-1) is found. Hence our analysis provides the necessary basis to study the wind measurement precision and spectral quality, which is prerequisite for reliable Eddy-Covariance flux measurements.
Metzger, S.; Junkermann, W.; Butterbach-Bahl, K.; Schmid, H. P.; Foken, T.
This study investigates whether the 3-D wind vector can be measured reliably from a highly transportable and low-cost weight-shift microlight aircraft. We draw up a transferable procedure to accommodate flow distortion originating from the aircraft body and -wing. This procedure consists of the analysis of aircraft dynamics and seven successive calibration steps. For our aircraft the horizontal wind components receive their greatest single amendment (14 %, relative to the initial uncertainty) from the correction of flow distortion magnitude in the dynamic pressure computation. Conversely the vertical wind component is most of all improved (31 %) by subsequent steps considering the 3-D flow distortion distribution in the flow angle computations. Therein the influences of the aircraft's trim (53 %), as well as changes in the aircraft lift (16 %) are considered by using the measured lift coefficient as explanatory variable. Three independent lines of analysis are used to evaluate the quality of the wind measurement: (a) A wind tunnel study in combination with the propagation of sensor uncertainties defines the systems input uncertainty to ≈0.6 m s-1 at the extremes of a 95 % confidence interval. (b) During severe vertical flight manoeuvres the deviation range of the vertical wind component does not exceed 0.3 m s-1. (c) The comparison with ground based wind measurements yields an overall operational uncertainty (root mean square error) of ≈0.4 m s-1 for the horizontal and ≈0.3 m s-1 for the vertical wind components. No conclusive dependence of the uncertainty on the wind magnitude (<8 m s-1) or true airspeed (ranging from 23-30 m s-1) is found. Hence our analysis provides the necessary basis to study the wind measurement precision and spectral quality, which is prerequisite for reliable Eddy-Covariance flux measurements.
Sridhar, Banavar (Inventor); Sheth, Kapil S. (Inventor); Chatterji, Gano Broto (Inventor); Bilimoria, Karl D. (Inventor); Grabbe, Shon (Inventor); Schipper, John F. (Inventor)
Methods for evaluating and implementing air traffic management tools and approaches for managing and avoiding an air traffic incident before the incident occurs. A first system receives parameters for flight plan configurations (e.g., initial fuel carried, flight route, flight route segments followed, flight altitude for a given flight route segment, aircraft velocity for each flight route segment, flight route ascent rate, flight route descent route, flight departure site, flight departure time, flight arrival time, flight destination site and/or alternate flight destination site), flight plan schedule, expected weather along each flight route segment, aircraft specifics, airspace (altitude) bounds for each flight route segment, navigational aids available. The invention provides flight plan routing and direct routing or wind optimal routing, using great circle navigation and spherical Earth geometry. The invention provides for aircraft dynamics effects, such as wind effects at each altitude, altitude changes, airspeed changes and aircraft turns to provide predictions of aircraft trajectory (and, optionally, aircraft fuel use). A second system provides several aviation applications using the first system. Several classes of potential incidents are analyzed and averted, by appropriate change en route of one or more parameters in the flight plan configuration, as provided by a conflict detection and resolution module and/or traffic flow management modules. These applications include conflict detection and resolution, miles-in trail or minutes-in-trail aircraft separation, flight arrival management, flight re-routing, weather prediction and analysis and interpolation of weather variables based upon sparse measurements. The invention combines these features to provide an aircraft monitoring system and an aircraft user system that interact and negotiate changes with each other.
Tholudin Mat Lazim
Full Text Available The main objective of the present work is to study the effect of an external store to a subsonic fighter aircraft. Generally most modern fighter aircraft is designed with an external store installation. In this project a subsonic fighter aircraft model has been manufactured using a computer numerical control machine for the purpose of studying the effect of the external store aerodynamic interference on the flow around the aircraft wing. A computational fluid dynamic (CFD and wind tunnel testing experiments have been carried out to ensure the aerodynamic characteristic of the model then certified the aircraft will not facing any difficulties in stability and controllability. In the CFD experiment, commercial CFD code is used to simulate the interference and aerodynamic characteristics of the model. Subsequently, the model together with an external store was tested in a low speed wind tunnel with test section sized 0.45 m×0.45 m. Result in the two-dimensional pressure distribution obtained by both experiments are comparable. There is only 12% deviation in pressure distribution found in wind tunnel testing compared to the result predicted by the CFD. The result shows that the effect of the external storage is only significant at the lower surface of the wing and almost negligible at the upper surface of the wing. Aerodynamic interference is due to the external storage were mostly evidence on a lower surface of the wing and almost negligible on the upper surface at low angle of attack. In addition, the area of influence on the wing surface by store interference increased as the airspeed increase.
Christhilf, David M.
It has long been recognized that frequency and phasing of structural modes in the presence of airflow play a fundamental role in the occurrence of flutter. Animation of simulation results for the long, slender Semi-Span Super-Sonic Transport (S4T) wind-tunnel model demonstrates that, for the case of mass-ballasted nacelles, the flutter mode can be described as a traveling wave propagating downstream. Such a characterization provides certain insights, such as (1) describing the means by which energy is transferred from the airflow to the structure, (2) identifying airspeed as an upper limit for speed of wave propagation, (3) providing an interpretation for a companion mode that coalesces in frequency with the flutter mode but becomes very well damped, (4) providing an explanation for bursts of response to uniform turbulence, and (5) providing an explanation for loss of low frequency (lead) phase margin with increases in dynamic pressure (at constant Mach number) for feedback systems that use sensors located upstream from active control surfaces. Results from simulation animation, simplified modeling, and wind-tunnel testing are presented for comparison. The simulation animation was generated using double time-integration in Simulink of vertical accelerometer signals distributed over wing and fuselage, along with time histories for actuated control surfaces. Crossing points for a zero-elevation reference plane were tracked along a network of lines connecting the accelerometer locations. Accelerometer signals were used in preference to modal displacement state variables in anticipation that the technique could be used to animate motion of the actual wind-tunnel model using data acquired during testing. Double integration of wind-tunnel accelerometer signals introduced severe drift even with removal of both position and rate biases such that the technique does not currently work. Using wind-tunnel data to drive a Kalman filter based upon fitting coefficients to
Merlaud, Alexis; Tack, Frederik; Constantin, Daniel; Fayt, Caroline; Maes, Jeroen; Mingireanu, Florin; Mocanu, Ionut; Georgescu, Lucian; Van Roozendael, Michel
The Small Whiskbroom Imager for atmospheric compositioN monitorinG (SWING) is an instrument dedicated to atmospheric trace gas retrieval from an Unmanned Aerial Vehicle (UAV). The payload is based on a compact visible spectrometer and a scanning mirror to collect scattered sunlight. Its weight, size, and power consumption are respectively 920 g, 27x12x12 cm3, and 6 W. The custom-built 2.5 m flying wing UAV is electrically powered, has a typical airspeed of 100 km/h, and can operate at a maximum altitude of 3 km. Both the payload and the UAV were developed in the framework of a collaboration between the Belgian Institute for Space Aeronomy (BIRA-IASB) and the Dunarea de Jos University of Galati, Romania. We present here SWING-UAV test flights dedicated to NO2 measurements and performed in Romania on 10 and 11 September 2014, during the Airborne ROmanian Measurements of Aerosols and Trace gases (AROMAT) campaign. The UAV performed 5 flights in the vicinity of the large thermal power station of Turceni (44.67° N, 23.4° E). The UAV was operated in visual range during the campaign, up to 900 m AGL , downwind of the plant and crossing its exhaust plume. The spectra recorded on flight are analyzed with the Differential Optical Absorption Spectroscopy (DOAS) method. The retrieved NO2 Differential Slant Column Densities (DSCDs) are up to 1.5e17 molec/cm2 and reveal the horizontal gradients around the plant. The DSCDs are converted to vertical columns and compared with coincident car-based DOAS measurements. We also present the near-future perspective of the SWING-UAV observation system, which includes flights in 2015 above the Black Sea to quantify ship emissions, the addition of SO2 as a target species, and autopilot flights at higher altitudes to cover a typical satellite pixel extent (10x10 km2).
Safi, Kamran; Kranstauber, Bart; Weinzierl, Rolf P.; Griffin, Larry; Reese, Eileen C.; Cabot, David; Cruz, Sebastian; Proaño, Carolina; Takekawa, John Y.; Newman, Scott H.; Waldenström, Jonas; Bengtsson, Daniel; Kays, Roland; Wikelski, Martin; Bohrer, Gil
Background: Understanding how environmental conditions, especially wind, influence birds' flight speeds is a prerequisite for understanding many important aspects of bird flight, including optimal migration strategies, navigation, and compensation for wind drift. Recent developments in tracking technology and the increased availability of data on large-scale weather patterns have made it possible to use path annotation to link the location of animals to environmental conditions such as wind speed and direction. However, there are various measures available for describing not only wind conditions but also the bird's flight direction and ground speed, and it is unclear which is best for determining the amount of wind support (the length of the wind vector in a bird’s flight direction) and the influence of cross-winds (the length of the wind vector perpendicular to a bird’s direction) throughout a bird's journey. Results: We compared relationships between cross-wind, wind support and bird movements, using path annotation derived from two different global weather reanalysis datasets and three different measures of direction and speed calculation for 288 individuals of nine bird species. Wind was a strong predictor of bird ground speed, explaining 10-66% of the variance, depending on species. Models using data from different weather sources gave qualitatively similar results; however, determining flight direction and speed from successive locations, even at short (15 min intervals), was inferior to using instantaneous GPS-based measures of speed and direction. Use of successive location data significantly underestimated the birds' ground and airspeed, and also resulted in mistaken associations between cross-winds, wind support, and their interactive effects, in relation to the birds' onward flight. Conclusions: Wind has strong effects on bird flight, and combining GPS technology with path annotation of weather variables allows us to quantify these effects for
Tholudin Mat Lazim
Full Text Available The main objective of the present work is to study the effect of an external store on a subsonic fighter aircraft. Generally most modern fighter aircrafts are designed with an external store installation. In this study, a subsonic fighter aircraft model has been manufactured using a computer numerical control machine for the purpose of studying the effect of the aerodynamic interference of the external store on the flow around the aircraft wing. A computational fluid dynamic (CFD simulation was also carried out on the same configuration. Both the CFD and the wind tunnel testing were carried out at a Reynolds number 1.86×105 to ensure that the aerodynamic characteristic can certify that the aircraft will not be face any difficulties in its stability and controllability. Both the experiments and the simulation were carried out at the same Reynolds number in order to verify each other. In the CFD simulation, a commercial CFD code was used to simulate the interference and aerodynamic characteristics of the model. Subsequently, the model together with an external store was tested in a low speed wind tunnel with a test section sized 0.45 m×0.45 m. Measured and computed results for the two-dimensional pressure distribution were satisfactorily comparable. There is only a 19% deviation between pressure distribution measured in wind tunnel testing and the result predicted by the CFD. The result shows that the effect of the external storage is only significant on the lower surface of the wing and almost negligible on the upper surface of the wing. Aerodynamic interference due to the external store was most evident on the lower surface of the wing and almost negligible on the upper surface at a low angle of attack. In addition, the area of influence on the wing surface by the store interference increased as the airspeed increased.
Lin, Raymond Chao
The handling qualities evaluation of nonlinear aircraft systems is an area of concern in loss-of-control (LOC) prevention. The Get Transfer Function (GetTF) method was demonstrated for evaluating the handling qualities of flight control systems and aircraft containing nonlinearities. NASA's Generic Transport Model (GTM), a nonlinear model of a civilian jet transport aircraft, was evaluated. Using classical techniques, the stability, control, and augmentation (SCAS) systems were designed to control pitch rate, roll rate, and airspeed. Hess's structural pilot model was used to model pilot dynamics in pitch and roll-attitude tracking. The simulated task was simultaneous tracking of, both, pitch and roll attitudes. Eight cases were evaluated: 1) gain increase of pitch-attitude command signal, 2) gain increase of roll-attitude command signal, 3) gain reduction of elevator command signal, 4) backlash in elevator actuator, 5) combination 3 and 4 in elevator actuator, 6) gain reduction of aileron command signal, 7) backlash in aileron actuator, and 8) combination of 6 and 7 in aileron actuator. The GetTF method was used to estimate the transfer function approximating a linear relationship between the proprioceptive signal of the pilot model and the command input. The transfer function was then used to predict the handling qualities ratings (HQR) and pilot-induced oscillation ratings (PIOR). The HQR is based on the Cooper-Harper rating scale. In pitch-attitude tracking, the nominal aircraft is predicted to have Level 2* HQRpitch and 2 pitch tracking exercise was also conducted to validate the structural pilot model.
赫瑞元; 马志微; 常丽萍; 郑仙荣
采用加压水热浸渍法制备氧化锌精脱硫剂,考察了浸渍时间、煅烧温度和煅烧时间对脱硫剂硫化性能的影响。对制备的脱硫剂在硫化温度500℃,空速2 000 h-1,气氛为H2（体积分数39%）、CO（体积分数33%）、H2S（体积分数1×10-3）和N2的条件下进行了固定床活性评价。结果显示,随着浸渍时间的延长,脱硫剂的脱硫能力增强,但硫化后脱硫剂粉化程度加剧。加压浸渍时间5 h、500℃煅烧5 h制备出的氧化锌脱硫剂的脱硫效果最佳。SEM结果表明,该条件下制备出的脱硫剂活性组分ZnO在载体表面分布均匀,且粒径较小。%Zinc oxide fine desulfurizer agent as the active constituent was prepared by press water hot dipping process.The influences of soaking time,calcination heat and calcination time on curing performance of the desulfurizer agent were studied.At the curing temperature of 500 ℃,airspeed of 2000 h-1,and H2（39 vol %）,CO（33 vol %）,H2S（1 000 ppm） or N2 as the atmosphere,activeness of the prepared desulfurizer agent on the fixed bed was appraised.The results demonstrated that along with soaking time expanding,desulfurizer agent desulphurization ability was strengthened,but the pulverization degree of cured desulfurizer agent was aggravated.The desulphurization effect of zinc oxide desulfurizer agent prepared in the condition of soaking 5 hours and calcining 5 hours at 500 ℃ was best.As the SEM result indicated,ZnO active constituent of the desulfurizer agent prepared in this condition was distributed in the carrier surface equably and the particle size was small.
A large planar microphone array,which consists of 111 microphones,was successfully applied to obtain a two-dimensional mapping of the sound sources on a landing aircraft.The focus of study in this paper is on the landing gear noise source.The spectra,directivities and sound pressure levels of flap side-edge noise of 7 narrow-board commercial aircraft and 7 wide-board commercial aircraft are presented.It is found that the landing gear noise spectrum is broadband with some single tones in some cases.The directivity of the total sound pressure level of a landing gear noise resembles that of a horizontal dipole.The level differences between the various aircraft landing gears are larger than those expected from the airspeed differences.It is thus expected that the louder noise emission of the landing gears can be reduced by redesigning.%应用由111个传声器组成的平面传声器阵列对当前流行的民用客机进场着陆过程中的机体噪声源进行了实验测量，本文对七架窄体客机和七架宽体客机的起落架噪声进行了分析，得到了起落架噪声的频谱特性、指向特性和声级变化.研究发现，起落架噪声的频谱是由宽频随机噪声与一些较为明显的单音噪声源组成，起落架噪声的指向性类似于一个水平放置的偶极子.不同飞机起落架噪声的声级相差较大，这说明可以通过重新结构设计降低起落架噪声。
other although they were measured in adjoining regions. The cirrus crystals in the maritime continental tropical region over Malaysia form tri-modal spectra that are not found in any of the other regions measured by the IAGOS aircraft so far, a feature that is possibly linked to biomass burning or dust. Frequent measurements of ice crystal concentrations greater than 1×105 L−1, often accompanied by anomalously warm temperature and erratic airspeed readings, suggest that aircraft often experience conditions that affect their sensors. This new instrument, if used operationally, has the potential of providing real-time and valuable information to assist in flight operations as well as providing real-time information for along-track nowcasting.
Full Text Available This study investigates whether the 3-D wind vector can be measured reliably from a highly transportable and low-cost weight-shift microlight aircraft. Therefore we draw up a transferable procedure to accommodate flow distortion originating from the aircraft body and -wing. This procedure consists of the analysis of aircraft dynamics and seven successive calibration steps. For our aircraft the horizontal wind components receive their greatest single amendment (14%, relative to the initial uncertainty from the correction of flow distortion magnitude in the dynamic pressure computation. Conversely the vertical wind component is most of all improved (31% by subsequent steps considering the 3-D flow distortion distribution in the flow angle computations. Therein the influences of the aircraft's aeroelastic wing (53%, as well as sudden changes in wing loading (16% are considered by using the measured lift coefficient as explanatory variable. Three independent lines of analysis are used to evaluate the quality of the wind measurement: (a A wind tunnel study in combination with the propagation of sensor uncertainties defines the systems input uncertainty to ≈0.6 m s−1 at the extremes of a 95% confidence interval. (b During severe vertical flight manoeuvres the deviation range of the vertical wind component does not exceed 0.3 m s−1. (c The comparison with ground based wind measurements yields an overall operational uncertainty (root mean square deviation of ≈0.4 m s−1 for the horizontal and ≈0.3 m s−1 for the vertical wind components. No conclusive dependence of the uncertainty on the wind magnitude (<8 m s−1 or true airspeed (ranging from 23–30 m s−1 is found. Hence our analysis provides the necessary basis to study the wind measurement precision and spectral quality, which is prerequisite for reliable eddy-covariance flux measurements.
Full Text Available This study investigates whether the 3-D wind vector can be measured reliably from a highly transportable and low-cost weight-shift microlight aircraft. Therefore we draw up a transferable procedure to accommodate flow distortion originating from the aircraft body and -wing. This procedure consists of the analysis of aircraft dynamics and seven successive calibration steps. For our aircraft the horizontal wind components receive their greatest single amendment (14 %, relative to the initial uncertainty from the correction of flow distortion magnitude in the dynamic pressure computation. Conversely the vertical wind component is most of all improved (31 % by subsequent steps considering the 3-D flow distortion distribution in the flow angle computations. Therein the influences of the aircraft's trim (53 %, as well as changes in the aircraft lift (16 % are considered by using the measured lift coefficient as explanatory variable. Three independent lines of analysis are used to evaluate the quality of the wind measurement: (a A wind tunnel study in combination with the propagation of sensor uncertainties defines the systems input uncertainty to ≈0.6 m s−1 at the extremes of a 95 % confidence interval. (b During severe vertical flight manoeuvres the deviation range of the vertical wind component does not exceed 0.3 m s−1. (c The comparison with ground based wind measurements yields an overall operational uncertainty (root mean square error of ≈0.4 m s−1 for the horizontal and ≈0.3 m s−1 for the vertical wind components. No conclusive dependence of the uncertainty on the wind magnitude (<8 m s−1 or true airspeed (ranging from 23–30 m s−1 is found. Hence our analysis provides the necessary basis to study the wind measurement precision and spectral quality, which is prerequisite for reliable Eddy-Covariance flux measurements.
R K Mehra; P O Arambel; A M Sampath; R K Prasanth; T C Parham
New algorithms and results are presented for flutter testing and adaptive notching of structural modes in V-22 tiltrotor aircraft based on simulated and flight-test data from Bell Helicopter Textron, Inc. (BHTI). For flutter testing and the identification of structural mode frequencies, dampings and mode shapes, time domain state space techniques based on Deterministic Stochastic Realization Algorithms (DSRA) are used to accurately identify multiple modessimultaneously from sine sweep and other multifrequency data, resulting in great savings over the conventional Prony method. Two different techniques for adaptive notching are explored in order to design an Integrated Flight Structural Control (IFSC) system. The first technique is based on on-line identification of structural mode parameters using DSRA algorithm and tuning of a notch filter. The second technique is based on decoupling rigid-body and structural modes of the aircraft by means of a Kalman filter and using rigid-body estimates in the feedback control loop. The difference between the two approaches is that on-line identification and adaptive notching in the first approach are entirely based on the knowledge of structural modes, whereas the Kalman filter design in the second approach is based on the rigid-body dynamic model only.In the first IFSC design, on-line identification is necessary for flight envelope expansion and to adjust the notch filter frequencies and suppress aero-servoelastic instabilities due to changing flight conditionssuch as gross weight, sling loads, and airspeed. It isshown that by tuning the notch filterfrequency to the identified frequency, the phase lag is reduced and the corresponding structural mode is effectively suppressed and stability is maintained. In the second IFSC design using Kalman filter design, the structural modes are again effectively suppressed. Furthermore, the rigid-body estimates are found to be fairly insensitive to both natural frequency and damping factor
神宁烯烃公司甲醇制丙烯装置，是目前世界上首套工业化甲醇制丙烯装置，甲醇首先经DME反应器转化成DME，之后进入甲醇制丙烯反应器转化为富含丙烯气的气烃化合物。DME反应器为绝热式固定床反应器，是目前国内最大的DME反应器，由于催化剂装填量大，床层为固定床，控温手段主要靠自身物料将反应热带走，在开车投料时极易发生飞温，造成催化剂失活。本文主要针对DME反应器在开车期间空速低，大量反应器热不能及时带走易造成飞温的情况提供操作指导。%Methanol to propylene device in Alkenes Company of Coal Chemical Industry Company Branch, Shenhua Ningxia Coal Indus-try Group is the first methanol to propylene device at present in the world. At first, methanol by DEM reactor into DME and then by methanol to propylene reactor into gas hydrocarbon compounds rich in propylene gas. DME reactor as the adiabatic fixed bed reactor is currently the largest DME reactor inland. Because of the large catalyst loading quantity and the fixed bed, as well as the means of tem-perature control mainly rely on their own materials to take away heat of reaction, the temperature runaway easily during their drive feed-ing leading to catalyst deactivation. This paper mainly for operating instructions to temperature runaway as DME reactor is low during driving airspeed and a lot of heat is not taken in time.
Conaway, Cody R.
From 2001-2011, the General Aviation (GA) fatal accident rate remained unchanged (Duquette & Dorr, 2014) with an overall stagnant accident rate between 2004 and 2013. The leading cause, loss of control in flight (NTSB, 2015b & 2015c) due to pilot inability to recognize approach to stall/spin conditions (NTSB, 2015b & 2016b). In 2013, there were 1,224 GA accidents in the U.S., accounting for 94% of all U.S. aviation accidents and 90% of all U.S. aviation fatalities that year (NTSB, 2015c). Aviation entails multiple challenges for pilots related to task management, procedural errors, perceptual distortions, and cognitive discrepancies. While machine errors in airplanes have continued to decrease over the years, human error still has not (NTSB, 2013). A preliminary analysis of a PC-based, Garmin G1000 flight deck was conducted with 3 professional pilots. Analyses revealed increased task load, opportunities for distraction, confusing perceptual ques, and hindered cognitive performance. Complex usage problems were deeply ingrained in the functionality of the system, forcing pilots to use fallible work arounds, add unnecessary steps, and memorize knob turns or button pushes. Modern computing now has the potential to free GA cockpit designs from knobs, soft keys, or limited display options. Dynamic digital displays might include changes in instrumentation or menu structuring depending on the phase of flight. Airspeed indicators could increase in size to become more salient during landing, simultaneously highlighting pitch angle on Attitude Indicators and automatically decluttering unnecessary information for landing. Likewise, Angle-of-Attack indicators demonstrate a great safety and performance advantage for pilots (Duquette & Dorr, 2014; NTSB, 2015b & 2016b), an instrument typically found in military platforms and now the Icon A5, light-sport aircraft (Icon, 2016). How does the design of pilots' environment---the cockpit---further influence their efficiency and
吴现印; 吕志果; 张伟伟
The new synthesis process of DL-pantolactone was researched, which employed glyoxalic acid and iso-butyraldehyde (IBD) as raw materials. The process included aldol condensation,catalytic hydrogenation of aldol condensate and dehydration-cycli-zation under nano copper-based catalyst in the fixed bed reactor. The results showed that the conservation of IBD was more than 95% under the following conditions in aldol condensation: reaction temperature 60 ℃, n (glyoxalic acid) : n (IBD) : n (TEA) = 1 : 1: 1. 2 and reaction time 2 h. Under the conditions of reaction temperature 120 *C, n (H2) : n(IBD) = 80 : 1, H2pressure 8. 0 Mpa, liquid airspeed 0. 6 ~ 1. 0 h-1, and vacuum distillation (the vacuity was 0.096 Mpa), the yield of DL-pantolactone was 90% , meanwhile the purity of the product reached 99. 65%.%研究了乙醛酸和异丁醛(IBD)在胺类催化剂作用下羟醛缩合及其缩合产物在纳米铜基催化剂上一步加氢脱水环化制备α-羟基-β,β-二甲基-γ-丁内酯的绿色合成新工艺.结果表明:在物料配比n(乙醛酸):n(异丁醛)∶n(三乙胺)为1∶1∶1.2,缩合温度60℃、反应时间2h的条件下羟醛缩合,IBD的转化率大于95％；将缩合产物采用负载于硅胶上的纳米铜基催化剂在120℃、n(H2)∶n(IBD)=80:1,氢气压力8.0 MPa、液空速0.6～1.0 h-1的条件下连续固定床加氢、环化脱水,减压蒸馏(真空度0.096 MPa),收集110 ～120℃馏分,DL-泛内酯的收率大于90％,产品纯度达到99.65％.
Chaney, Christopher Scott
A study was conducted to assess the accuracy of empirical techniques used for the calculation of flight performance for unmanned aerial vehicles. This was achieved by quantifying the error between a mathematical model developed with these techniques and experimental test data taken using an unmanned aircraft. The vehicle utilized for this study was developed at Washington State University for the purpose of flying using power derived from hydrogen stored as a cryogenic liquid. The vehicle has a mass of 32.8 kg loaded and performed a total of 14 flights under battery power for 3.58 total flight hours. Over these flights, the design proved it is capable of sustaining level flight from the power available from a PEM fuel cell propulsion system. The empirical techniques used by the model are explicitly outlined within. These yield several performance metrics that are compared to measurements taken during flight testing. Calculations of required thrust for steady flight over all airspeeds and rates of climb modeled are found to have a mean percent error of 3.2%+/-7.0% and a mean absolute percent error of 34.6%+/-5.1%. Comparison of the calculated and measured takeoff distance are made and the calculated thrust required to perform a level turn at a given rate is compared to flight test data. A section of a test flight is analyzed, over which the vehicle proves it can sustain level flight under 875 watts of electrical power. The aircraft's design is presented including the wing and tail, propulsion system, and build technique. The software and equipment used for the collection and analysis of flight data are given. Documentation and validation is provided of a unique test rig for the characterization of propeller performance using a car. The aircraft remains operational to assist with research of alternative energy propulsion systems and novel fuel storage techniques. The results from the comparison of the mathematical model and flight test data can be utilized to assist
Wing, David J.; Smith, Jeremy C.; Ballin, Mark G.
The Dynamic Multi-track Airways (DMA) Concept for Air Traffic Management (ATM) proposes a network of high-altitude airways constructed of multiple, closely spaced, parallel tracks designed to increase en-route capacity in high-demand airspace corridors. Segregated from non-airway operations, these multi-track airways establish high-priority traffic flow corridors along optimal routes between major terminal areas throughout the National Airspace System (NAS). Air traffic controllers transition aircraft equipped for DMA operations to DMA entry points, the aircraft use autonomous control of airspeed to fly the continuous-airspace airway and achieve an economic benefit, and controllers then transition the aircraft from the DMA exit to the terminal area. Aircraft authority within the DMA includes responsibility for spacing and/or separation from other DMA aircraft. The DMA controller is responsible for coordinating the entry and exit of traffic to and from the DMA and for traffic flow management (TFM), including adjusting DMA routing on a daily basis to account for predicted weather and wind patterns and re-routing DMAs in real time to accommodate unpredicted weather changes. However, the DMA controller is not responsible for monitoring the DMA for traffic separation. This report defines the mature state concept, explores its feasibility and performance, and identifies potential benefits. The report also discusses (a) an analysis of a single DMA, which was modeled within the NAS to assess capacity and determine the impact of a single DMA on regional sector loads and conflict potential; (b) a demand analysis, which was conducted to determine likely city-pair candidates for a nationwide DMA network and to determine the expected demand fraction; (c) two track configurations, which were modeled and analyzed for their operational characteristic; (d) software-prototype airborne capabilities developed for DMA operations research; (e) a feasibility analysis of key attributes in
Bange, J.; van den Kroonenberg, A. C.; Spieß, T.; Buschmann, M.; Krüger, L.; Heindorf, A.; Vörsmann, P.
The limitations of manned airborne meteorological measurements led to the development of an autonomously operating mini aircraft, the Meteorological Mini-UAV (M2AV), at the Institute of Aerospace Systems, Technical University of Braunschweig, Germany. The task was to develop, test and verify a meteorological sensor package as payload for an already available automatic carrier aircraft, the UAV 'Carolo T200', which operates autonomously i.e. without remote control. The M2AV is a self constructed model aircraft with two electrically powered engines and a wingspan of two meters. The maximum take-off weight is 4.5~kg (the M2AV is therefore classified as an model plane which simplifies authority issues), including 1.5~kg of payload. It is hand-launched which makes operation of the aircraft easy. With an endurance of approximately 50 minutes, the range accounts for 60 km at a cruising speed of 20~m/s. The M2AV is capable of performing turbulence measurements (wind vector, temperature and humidity) within the troposphere and offers an economic component during meteorological campaigns. The meteorological sensors are mounted on a noseboom to minimise the aircraft's influence on the measurements and to position the sensors closely to each other. Wind is measured via a small five-hole probe, an inertia measurement unit and GPS. The flight mission (waypoints, altitudes, airspeed) is planned and assigned to the aircraft before the semi- automatic launch. The flight is only controlled by the on-board autopilot system which only communicates with a ground station (laptop PC) for the exchange of measured data and command updates like new waypoints etc. The talk gives details on the technical items, calibration and first missions. Results from first field experiments like the LAUNCH-2005 campaign near Berlin are used for data quality assessment by comparison with simultaneous lidar and sodar measurements. An in situ comparison with the highly accurate helicopter-borne turbulence
Full Text Available In this paper, we show that in mixed phase clouds, the presence of ice crystals may induce wrong FSSP 100 measurements interpretation especially in terms of particle size and subsequent bulk parameters. The presence of ice crystals is generally revealed by a bimodal feature of the particle size distribution (PSD. The combined measurements of the FSSP-100 and the Polar Nephelometer give a coherent description of the effect of the ice crystals on the FSSP-100 response. The FSSP-100 particle size distributions are characterized by a bimodal shape with a second mode peaked between 25 and 35 μm related to ice crystals. This feature is observed with the FSSP-100 at airspeed up to 200 m s−1 and with the FSSP-300 series. In order to assess the size calibration for clouds of ice crystals the response of the FSSP-100 probe has been numerically simulated using a light scattering model of randomly oriented hexagonal ice particles and assuming both smooth and rough crystal surfaces. The results suggest that the second mode, measured between 25 μm and 35 μm, does not necessarily represent true size responses but corresponds to bigger aspherical ice particles. According to simulation results, the sizing understatement would be neglected in the rough case but would be significant with the smooth case. Qualitatively, the Polar Nephelometer phase function suggests that the rough case is the more suitable to describe real crystals. Quantitatively, however, it is difficult to conclude. A review is made to explore different hypotheses explaining the occurrence of the second mode. However, previous cloud in situ measurements suggest that the FSSP-100 secondary mode, peaked in the range 25–35 μm, is likely to be due to the shattering of large ice crystals on the probe inlet. This finding is supported by the rather good relationship between the concentration of particles larger than 20 μm (hypothesized to be ice shattered-fragments measured by the
郭军; 董新民; 王龙
To keep the position of an unmanned combat air vehicle （UCAV） during autonomous aerial refueling, we develop the dynamic model and put forward a nonlinear controller for the UCAV with time-varying mass. By comprehen- sively considering the effect of fuel transfer on the UCAV mass, the inertia matrix and the center of mass, we derive the time-varying mass dynamic equations of UCAV based on state variables relative to inertial reference frame. By introducing the spectral radius, we apply the command filtered backstepping （CFBS） method based on localized adaptive bounds to control the UCAV in a desired position. The unknown model uncertainties are approximated online by using approxima- tors. Localized adaptive bounds are used to compensate inherent approximation errors and external disturbances. Using CFBS, we design four feedback control loops for the relative position, airspeed, attitude angle, and angular rate to guarantee the stability of the UCAV. Nonlinear simulation demonstrates the effectiveness of the nonlinear flight control law in three different refueling cases.%针对自主空中加油中无人作战飞机（UCAV）位置保持问题，进行了时变质量UCAV的动力学建模与非线性控制设计．综合考虑了燃油传输对UCAV的质量、惯性矩阵和质心位置的影响，基于相对于惯性系的状态变量，推导了uCAV时变质量动力学方程．通过引入谱半径，将局部化自适应边界指令滤波反推方法应用于UCAV的位置保持控制．使用逼近器对未知模型不确定性进行在线逼近．对于固有逼近误差和外部扰动，采用局部化自适应边界进行补偿．通过指令滤波反推，设计了相对位置、速度、姿态角和角速度四个反馈回路来保证UCAV的稳定性．最后，三种不同加油方案下的非线性仿真验证了非线性飞行控制律的有效性．
李荣冰; 于永军; 刘建业; 熊智
高精度导航系统作为无人机自主控制的核心,已成为制约无人机性能提升的关键因素.SINS/GPS( strapdown inertial navigation system/global positioning system)组合导航系统由于GPS信号易受干扰而无法满足无人机长时间稳定精确导航的需要.在分析大气数据系统特性的基础上,推导了捷联惯导与大气空速、高度误差模型,建立了SINS/ADS( air data system)量测方程,并设计了大气数据系统辅助的SINS/GPS松组合导航系统；针对ADS和GPS输出频率不一致的问题,提出了时间更新和量测更新分离的异步滤波算法.实际跑车验证结果表明,设计的系统能够在GPS受干扰的情况下有效提高系统的精度和稳定性,定位精度可以达到25 m.%High precision navigation system is the key technology of UAV ( unmanned air vehicles) autonomous control ,which has been a major constraint on the improvement of UAV performance. SINS/GPS ( strapdown inertial navigation system/global positioning system) integrated navigation system can not meet the need of UAV navigation accuracy and stability in long time because GPS signal is susceptible to interference. Based on the analysis of the characteristics of air data system,the SINS/ADS (air data system) error model is derived based on the attitude and airspeed information provided by pressure sensors,and the SINS/ADS measurement equation is established. The SINS/ GPS integrated navigation system assisted by air data system is designed. Aiming at the output frequency inconsistency between GPS and ADS,an asynchronous centralized Kalman filter is designed for separating the time update period and measurement update period. Actual operation experiments on vehicle confirm that the designed system can effectively improve the accuracy and stability of the system under interference condition,and the positioning accuracy can reach 25 m.
Buchholz, Bernhard; Ebert, Volker
Airborne hygrometry is often demanded in scientific flight campaigns to provide datasets for environmental modeling or to correct for water vapor dilution or cross sensitivity effects in other gas analytical techniques. Water vapor measurements, however, are quite challenging due to the large dynamic range in the atmosphere (between 2 and 40000 ppmv) and the high spatio-temporal variability. Airborne hygrometers therefore need to combine a large measurement range with high temporal resolution to resolve - at typical airspeeds of 500 to 900 km/h - atmospheric gradients of several 1000 ppmv/s. Especially during the ascent into the upper troposphere, hygrometers need to work at high gas exchange rates to minimize water vapor adsorption effects. On the other hand, water vapor sensors are difficult to calibrate due to the strong water adsorption and the lack of bottled reference gas standards, which requires pre- or/and post-flight field calibrations. Recently in-flight calibration using an airborne H2O generator was demonstrated, which minimizes calibration drift but still imposes a lot of additional work and hardware to the experiments, since these kind of calibrations just transfer the accuracy level issues to the in-flight calibration-source. To make things worse, the low gas flow (1-5 std l/min, compared with up to 100 std l/min in flight for fast response instruments) adheres critical questions of wall absorption/desorption of the source and instrument even during the calibration process. The national metrological institutes (NMIs) maintain a global metrological water vapor scale which is defined via national primary humidity generators. These provide for calibration purposes well-defined, accurate water vapor samples of excellent comparability and stability traced back to the SI-units. The humidity calibration chain is maintained via high accuracy (but rather slow) Dew-Point-Mirror-Hygrometers as transfer standards. These provide a traceable performance and
Peterson, Randall L.; Warmbrodt, William (Technical Monitor)
The tilt rotor aircraft holds great promise for improving air travel in the future. It's benefits include vertical take off and landing combined with airspeeds comparable to propeller driven aircraft. However, the noise from a tilt rotor during approach to a landing is potentially a significant barrier to widespread acceptance of these aircraft. This approach noise is primarily caused by Blade Vortex Interactions (BVI), which are created when the blade passes near or through the vortex trailed by preceding blades. The XV- 15 Aeroacoustic test will measure the noise from a tilt rotor during descent conditions and demonstrate several possible techniques to reduce the noise. The XV- 15 Aeroacoustic test at NASA Ames Research Center will measure acoustics and performance for a full-scale XV-15 rotor. A single XV-15 rotor will be mounted on the Ames Rotor Test Apparatus (RTA) in the 80- by 120-Foot Wind Tunnel. The test will be conducted in helicopter mode with forward flight speeds up to 100 knots and tip path plane angles up to +/- 15 degrees. These operating conditions correspond to a wide range of tilt rotor descent and transition to forward flight cases. Rotor performance measurements will be made with the RTA rotor balance, while acoustic measurements will be made using an acoustic traverse and four fixed microphones. The acoustic traverse will provide limited directionality measurements on the advancing side of the rotor, where BVI noise is expected to be the highest. Baseline acoustics and performance measurements for the three-bladed rotor will be obtained over the entire test envelope. Acoustic measurements will also be obtained for correlation with the XV-15 aircraft Inflight Rotor Aeroacoustic Program (IRAP) recently conducted by Ames. Several techniques will be studied in an attempt to reduce the highest measured BVI noise conditions. The first of these techniques will use sub-wings mounted on the blade tips. These subwings are expected to alter the size
Delay banking has been invented to enhance air-traffic management in a way that would increase the degree of fairness in assigning arrival, departure, and en-route delays and trajectory deviations to aircraft impacted by congestion in the national airspace system. In delay banking, an aircraft operator (airline, military, general aviation, etc.) would be assigned a numerical credit when any of their flights are delayed because of an air-traffic flow restriction. The operator could subsequently bid against other operators competing for access to congested airspace to utilize part or all of its accumulated credit. Operators utilize credits to obtain higher priority for the same flight, or other flights operating at the same time, or later, in the same airspace, or elsewhere. Operators could also trade delay credits, according to market rules that would be determined by stakeholders in the national airspace system. Delay banking would be administered by an independent third party who would use delay banking automation to continually monitor flights, allocate delay credits, maintain accounts of delay credits for participating airlines, mediate bidding and the consumption of credits of winning bidders, analyze potential transfers of credits within and between operators, implement accepted transfers, and ensure fair treatment of all participating operators. A flow restriction can manifest itself in the form of a delay in assigned takeoff time, a reduction in assigned airspeed, a change in the position for the aircraft in a queue of all aircraft in a common stream of traffic (e.g., similar route), a change in the planned altitude profile for an aircraft, or change in the planned route for the aircraft. Flow restrictions are typically imposed to mitigate traffic congestion at an airport or in a region of airspace, particularly congestion due to inclement weather, or the unavailability of a runway or region of airspace. A delay credit would be allocated to an operator of a
吕维阳; 刘盛余; 能子礼超; 汪雪婷; 吴萧
在固定床反应装置上考察了商业载硫活性炭脱除天然气中气态 Hg0的吸附性能、影响因素、再生方法以及吸附动力学,同时结合BET,FTIR,XRD等表征手段提出载硫活性炭脱汞机理.结果表明,空速对脱汞效率限制作用有限,空速从12000h-1提高至48000h-1,脱除率变化范围在 7%以内.增加汞浓度在初始阶段可以提高其脱汞率,增加单位质量活性炭对汞的吸附量,提高温度会增加吸附体系内的活化分子,提高脱汞率,温度在80℃时效果最优,但是温度过高则会产生负面效应.不同阶段的动力学拟合结果表明化学吸附是整个吸附过程的控制步骤.热脱附实验表明载硫活性炭的脱汞温度是在 300~450℃,再生后对汞的吸附能力减弱,其原因可归为再生过程中碳硫键的损失和活性炭二次碳化时表面的烧蚀.%The adsorption performance, influence factors, regeneration method and the adsorption dynamics of the gaseous Hg0 in natural gas removed by commercial sulfur-loaded activated carbon were studied on fixed reacting device. Meanwhile, the removal mercury mechanism of S-loaded activated carbon was put forward by applying characterization methods, including BET, FTIR, XRD and etc. The result showed that the restriction effect of airspeed on removal mercury efficiency was limited. In the initial stage, removal mercury rate and adsorption capacity of element mercury can be raised by increasing mercury concentration. As temperature gets higher, the activated molecule in the system get higher and the removal mercury rate will be promoted. The best effect appeared at 80℃. But when temperature was higher than 80℃, negative effects arose. Dynamic fitting in different stages showed that chemical adsorption was the controlling step of the whole process. Besides, the thermal desorption experiment showed that the removal mercury temperature of s-loaded activated carbon ranged from 300℃ to 450℃, and the
F-T synthetic oil is mainly composed of n-paraffins of alkanes and alkenes,in which wax is accounted for 40%. The production maximization of the high quality middle distillates are one of the effec-tive ways to make full use of F-T synthetic wax. The development and application of the catalysts with high performance is the key of hydrocracking technology. The recent research progress in hydrocracking cata-lysts and their application status at home and abroad were reviewed. The commercial applications of FC-14 and SC-I catalysts in Inner Mongolia Yitai CTO Co. ,Ltd. were especially introduced. The running results showed that SC-I catalyst exhibited good activity and selectivity to middle distillate. Under the condition of total feedstock rate 20 t·h-1,hydrogen partial pressure of reactor inlet 7. 0 MPa,hydrogen/oil volume ratio 800 ,total volume airspeed( 2 . 0-15 . 0 )h-1 ,reactor outlet temperature about 340 ℃ and overall temperature rise 14 ℃,the catalyst exhibited high reaction activity,flexible temperature control perform-ance and high selectivity to medium oil,and diesel yield increased about 3 -4 percentage point at the same time.%F-T合成油低温工艺产物中有大于质量分数40%的石蜡生成，必须对其进行加氢裂化生产优质的中间馏分油。加氢裂化关键在于高性能催化剂的研究开发，概述了近年来加氢裂化催化剂在国内外的应用现状。FC-14及SC-I催化剂在内蒙古伊泰煤制油有限责任公司的运行结果表明，SC-I催化剂表现出良好的活性及较高的中间馏分油选择性，在总进料20 t·h-1、反应器入口氢分压7.0 MPa、氢油体积比800和总体积空速（2.0~15.0）h-1条件下，反应器出口温度约340℃，总温升14℃，表现出较高的反应活性及灵活的温度调控性，柴油收率上升约3~4个百分点，具有较高的中油选择性。
Buchholz, Bernhard; Afchine, Armin; Krämer, Martina; Ebert, Volker
in a so called "closed-path" cell  for total water measurement via a forward facing inlet. The other part of the laser light is coupled to an "open-path" cell  placed outside of the aircraft fuselage to measure gas phase water without any possible artifacts from ice or liquid particles. The frequency of the measurements can be up to 240 Hz (4.2 msec) for all four channels. Altogether, the novel HAI instrument allows fast, accurate and precise dual-phase water measurements. The individual evaluation of the multi-channel raw-data is done afterwards, without any channel interdependencies, in a calibration-free mode. The water signals are combined with an extensive set of more than 100 housekeeping data to enable a holistic data quality management and a rigorous signal scrutiny to maximize the confidence level of the final H2O values. HAI therefore represents a new unique research tool for atmospheric hygrometry to address numerous open topics in atmospheric research. First scientific HAI campaigns have been successfully realized in 2012 onboard the German research plane HALO (High Altitude and Long Range Research Aircraft) during the TACTS and ESMVal missions. The first two HALO campaigns in clouds (MLCIRRUS and ACRIDICON) will be realized in 2014. In our contribution we present and discuss the performance of HAI and show detailed evaluations of typical inflight data. The results of the first two HAI campaigns on HALO resulted in more than 100 operation hours of continuous data and show nice agreement between the closed-path and open-path under clear sky conditions, despite the different sampling conditions of the sensor channels and airspeed of up to 900 km/h in the open path section. All mission data are and will be uploaded to the HALO database and are available for further scientific exploitation. Furthermore, the HAI principle can be adapted to other (airborne) platforms and be used for phase resolved science of the atmospheric water cycle. In parallel HAI
This 1992 photo shows SR-71 flight engineer Marta Bohn-Meyer in front of one of NASA's SR-71 aircraft on the ramp at the Ames-Dryden Flight Research Facility (later, Dryden Flight Research Center), Edwards, California. An aerospace engineer who has been at Dryden since 1979, Bohn-Meyer is the first female crew member ever assigned to fly in the SR-71. Data from the SR-71 program carried out by NASA will be used to aid designers of future supersonic aircraft and propulsion systems. Two SR-71 aircraft have been used by NASA as testbeds for high-speed and high-altitude aeronautical research. The aircraft, an SR-71A and an SR-71B pilot trainer aircraft, have been based here at NASA's Dryden Flight Research Center, Edwards, California. They were transferred to NASA after the U.S. Air Force program was cancelled. As research platforms, the aircraft can cruise at Mach 3 for more than one hour. For thermal experiments, this can produce heat soak temperatures of over 600 degrees Fahrenheit (F). This operating environment makes these aircraft excellent platforms to carry out research and experiments in a variety of areas -- aerodynamics, propulsion, structures, thermal protection materials, high-speed and high-temperature instrumentation, atmospheric studies, and sonic boom characterization. The SR-71 was used in a program to study ways of reducing sonic booms or over pressures that are heard on the ground, much like sharp thunderclaps, when an aircraft exceeds the speed of sound. Data from this Sonic Boom Mitigation Study could eventually lead to aircraft designs that would reduce the 'peak' overpressures of sonic booms and minimize the startling affect they produce on the ground. One of the first major experiments to be flown in the NASA SR-71 program was a laser air data collection system. It used laser light instead of air pressure to produce airspeed and attitude reference data, such as angle of attack and sideslip, which are normally obtained with small tubes and vanes
The two pilot-engineer teams that flew the SR-71 aircraft at the NASA Ames-Dryden Flight Research Facility (later, Dryden Flight Research Center, Edwards, California, are, from left, pilot Rogers Smith, flight engineers Robert Meyer and Marta Bohn-Meyer, and pilot Steven Ishmael. The Meyers are the first husband-wife team of aeronautical engineers at Dryden on flight status. Two SR-71 aircraft have been used by NASA as testbeds for high-speed and high-altitude aeronautical research. The aircraft, an SR-71A and an SR-71B pilot trainer aircraft, have been based here at NASA's Dryden Flight Research Center, Edwards, California. They were transferred to NASA after the U.S. Air Force program was cancelled. As research platforms, the aircraft can cruise at Mach 3 for more than one hour. For thermal experiments, this can produce heat soak temperatures of over 600 degrees Fahrenheit (F). This operating environment makes these aircraft excellent platforms to carry out research and experiments in a variety of areas -- aerodynamics, propulsion, structures, thermal protection materials, high-speed and high-temperature instrumentation, atmospheric studies, and sonic boom characterization. The SR-71 was used in a program to study ways of reducing sonic booms or over pressures that are heard on the ground, much like sharp thunderclaps, when an aircraft exceeds the speed of sound. Data from this Sonic Boom Mitigation Study could eventually lead to aircraft designs that would reduce the 'peak' overpressures of sonic booms and minimize the startling affect they produce on the ground. One of the first major experiments to be flown in the NASA SR-71 program was a laser air data collection system. It used laser light instead of air pressure to produce airspeed and attitude reference data, such as angle of attack and sideslip, which are normally obtained with small tubes and vanes extending into the airstream. One of Dryden's SR-71s was used for the Linear Aerospike Rocket Engine, or
longitudinal (pitch) motions; wing and tail loads, lift, drag, and buffeting characteristics of swept-wing aircraft at transonic and supersonic speeds; and the effects of the rocket exhaust plume on lateral dynamic stability throughout the speed range. (Plume effects were a new experience for aircraft.) The number three aircraft also gathered information about the effects of external stores (bomb shapes, drop tanks) upon the aircraft's behavior in the transonic region (roughly 0.7 to 1.3 times the speed of sound). In correlation with data from other early transonic research aircraft such as the XF-92A, this information contributed to solutions to the pitch-up problem in swept-wing aircraft. The three airplanes flew a total of 313 times--123 by the number one aircraft (Bureau No. 37973--NACA 143), 103 by the second Skyrocket (Bureau No. 37974--NACA 144), and 87 by airplane number three (Bureau No. 37975--NACA 145). Skyrocket 143 flew all but one of its missions as part of the Douglas contractor program to test the airplane's performance. NACA aircraft 143 was initially powered by a Westinghouse J-34-40 turbojet engine configured only for ground take-offs, but in 1954-55 the contractor modified it to an all-rocket air-launch capability featuring an LR8-RM-6, 4-chamber Reaction Motors engine rated at 6,000 pounds of thrust at sea level (the Navy designation for the Air Force's LR-11 used in the X-1). In this configuration, NACA research pilot John McKay flew the airplane only once for familiarization on September 17, 1956. The 123 flights of NACA 143 served to validate wind-tunnel predictions of the airplane's performance, except for the fact that the airplane experienced less drag above Mach 0.85 than the wind tunnels had indicated. NACA 144 also began its flight program with a turbojet powerplant. NACA pilots Robert A. Champine and John H. Griffith flew 21 times in this configuration to test airspeed calibrations and to research longitudinal and lateral stability and control
to test airspeed calibrations and to research longitudinal and lateral stability and control. In the process, during August of 1949 they encountered pitch-up problems, which NACA engineers recognized as serious because they could produce a limiting and dangerous restriction on flight performance. Hence, they determined to make a complete investigation of the problem. In 1950, Douglas replaced the turbojet with an LR-8 rocket engine, and its pilot, William B. Bridgeman, flew the aircraft seven times up to a speed of Mach 1.88 (1.88 times the speed of sound) and an altitude of 79,494 feet (the latter an unofficial world's altitude record at the time, achieved on August 15, 1951). In the rocket configuration, a Navy P2B (Navy version of the B-29) launched the airplane at approximately 30,000 feet after taking off from the ground with the Skyrocket attached beneath its bomb bay. During Bridgeman's supersonic flights, he encountered a violent rolling motion known as lateral instability that was less pronounced on the Mach 1.88 flight on August 7, 1951, than on a Mach 1.85 flight in June when he pushed over to a low angle of attack (angle of the fuselage or wing to the prevailing wind direction). The NACA engineers studied the behavior of the aircraft before beginning their own flight research in the airplane in September 1951. Over the next couple of years, NACA pilot A. Scott Crossfield flew the airplane 20 times to gather data on longitudinal and lateral stability and control, wing and tail loads, and lift, drag, and buffeting characteristics at speeds up to Mach 1.878. At that point, Marine Lt. Col. Marion Carl flew the airplane to a new (unofficial) altitude record of 83,235 feet on August 21, 1953, and to a maximum speed of Mach 1.728. Following Carl's completion of these flights for the Navy, NACA technicians at the High-Speed Flight Research Station (HSFRS) near Mojave, Calif., outfitted the LR-8 engine's cylinders with nozzle extensions to prevent the exhaust gas
. Champine and John H. Griffith flew 21 times in this configuration to test airspeed calibrations and to research longitudinal and lateral stability and control. In the process, during August of 1949 they encountered pitch-up problems, which NACA engineers recognized as serious because they could produce a limiting and dangerous restriction on flight performance. Hence, they determined to make a complete investigation of the problem. In 1950, Douglas replaced the turbojet with an LR-8 rocket engine, and its pilot, William B. Bridgeman, flew the aircraft seven times up to a speed of Mach 1.88 (1.88 times the speed of sound) and an altitude of 79,494 feet (the latter an unofficial world's altitude record at the time, achieved on August 15, 1951). In the rocket configuration, a Navy P2B (Navy version of the B-29) launched the airplane at approximately 30,000 feet after taking off from the ground with the Skyrocket attached beneath its bomb bay. During Bridgeman's supersonic flights, he encountered a violent rolling motion known as lateral instability that was less pronounced on the Mach 1.88 flight on August 7, 1951, than on a Mach 1.85 flight in June when he pushed over to a low angle of attack (angle of the fuselage or wing to the prevailing wind direction). The NACA engineers studied the behavior of the aircraft before beginning their own flight research in the airplane in September 1951. Over the next couple of years, NACA pilot A. Scott Crossfield flew the airplane 20 times to gather data on longitudinal and lateral stability and control, wing and tail loads, and lift, drag, and buffeting characteristics at speeds up to Mach 1.878. At that point, Marine Lt. Col. Marion Carl flew the airplane to a new (unofficial) altitude record of 83,235 feet on August 21, 1953, and to a maximum speed of Mach 1.728. Following Carl's completion of these flights for the Navy, NACA technicians at the High-Speed Flight Research Station (HSFRS) near Mojave, Calif., outfitted the LR-8 engine