Hovering Helicopter Research Articles

Efficient Minimax Control Design for Prescribed Parameter Uncertainty

The purview of this paper is the design of a controller that includes knowledge of parametric uncertainties and their distributions. The parameter distributions are approximated by a finite set of points that are calculated by the unscented transformation. This set of points is used to design robust controllers that minimize the worst performance of the plant over the domain of uncertainty. The proposed technique is illustrated on two benchmark problems. The first relates to the design of prefilters for a spring‐mass‐dashpot system and the second is the feedback control of a hovering helicopter. Numerical simulations are used to illustrate the proximity of the resulting controller to the minimax controller.

Visibility of St Lawrence belugas to aerial photography, estimated by direct observation

The depleted population of belugas (Delphinapterus leucas) inhabiting the St Lawrence estuary, Canada, was monitored by periodic photographic aerial surveys. In order to correct counts made on aerial survey film and to obtain an estimate of the true size of the population, the diving behaviour and the visibility from the air of these animals was studied. A Secchi-disk turbidity survey in the belugas’ summer range showed that water clarity varied between 1.5 m and 11.6 m. By studying aerial photographs of sheet-plastic models of belugas that had been sunk to different depths below the surface, we found that models of white adults could be seen down to about the same depth as a Secchi disk, but no deeper. Smaller models of dark-grey juveniles could only be seen down to about 50% of Secchi-disk depth. By observing groups of belugas from a hovering helicopter and recording their disappearances and re-appearances, it was found that they were visible for 44.3% of the time, and that an appropriate correction for single photographs would be to multiply the photographic count by about 222% (SE 20%). For surveys in which there was overlap between adjacent frames, the estimated correction would be 209% (SE 16%). This correction factor was slightly conservative and gave an estimate of the true size of the population, based on a single survey, of 1,202 belugas (SE 189) in 1997. An estimate for 1997 based on smoothing 5 surveys 1988–1997 was 1,238 (SE 119).

Helicopter flight around a ship's' superstructure

A computational fluid dynamics model of a hovering helicopter main rotor is developed to examine air flow in the presence of ship structures and side winds. An illustration of the problem is given. The rotor is modelled by modifying the governing Navier-Stokes equations in the region of the disc. The extra terms added to the governing equations apply a downward force to the fluid; these forces are independent of the flow around the rotor and are equal to the helicopter weight. The helicopter rotor model and the ship model are combined to yield one flow solution, which, due to the severe non-linearities of the problem, cannot be achieved by superposition. The resultant flow yields valuable data about the induced velocities at the rotor, which ultimately determine the control pitch and power required to maintain the hover in a given location. Indeed, the interactions between the rotor downwash and ship air flow are known to produce unexpected and adverse flight dynamic behaviour of the aircraft.

Performance prediction and flowfield analysis of rotors in hover, using a coupled Euler/boundary layer method

The performance prediction of helicopter in hover is of key importance for manufacturers because hover is a design configuration for the definition of a rotorcraft. A lot of effort has been made for more than 10 years all over the world in order to develop and validate numerical methods based on CFD. An Euler method (WAVES) developed by ONERA and coupled with a boundary layer code (MI3DI) is presented, validated and applied to compute the total performance of rotors with different tip shapes. A new boundary condition for the Euler code has been tested and enables better calculation by eliminating `numerical' recirculation. The code has demonstrated its ability to rank two rotors with different planforms in good agreement with experiment. Under industrial requirements new grid strategies have been studied and should allow to reduce CPU time consumption. It is shown that WAVES/MI3DI can be efficiently used in the aerodynamic design process of a new rotor.

The mt. tyndall incident

The authors describe the 53-hour rescue of a 6-foot, 11-inch tall, 250-pound hiker in the face of harsh environmental conditions in Sequoia National Park. This 43-year-old man fell 25 feet, injured his leg, and was noted to be hypothermic and hypovolemic. Weather, altitude, and the patient's size delayed and complicated his evacuation. After being carried down 1,500 vertical feet, he was hoisted into a hovering helicopter and flown to University Medical Center in Fresno, California. On arrival, the patient was determined to have a comminuted subtrochanteric right femur fracture, which was ultimately repaired surgically. The authors also discuss some of the unique aspects of wilderness and National Park Service EMS.

Identification of Linear Models for a Hovering Helicopter Rotor

Abstract Rotor control techniques for active control of vibrations require the availability of dynamic models of the response of the rotor to control inputs. The results of a simulation study of black-box identification trials aiming at the derivation of reduced order, discrete-time, linear models for the response of helicopter rotor loads to individual blade control inputs are presented and discussed. The study has been performed by making use of a simulation program for the dynamics and aerodynamics of the rotor of an existing aircraft, and three identification algorithms have been compared: a frequency domain method and two time domain, subspace based methods.

Linear quadratic optimal model-following control of a helicopter in hover

Presented are results of a study of the application of linear quadratic optimal model-following control applied to a Bell 205 helicopter in hover. The primary objective of good in-flight stability robustness and performance was accomplished via singular value analysis using perturbed systems. Nominal aircraft models were compared with experimental data and discrepancies quantified in a robustness criterion. Current military handling quality requirements were specified as a target model to be followed. The linear quadratic optimal control and command feedforward was found suitable for these requirements. Design analyses enabled consideration of the tuning process, where effects of variations in selected tuning parameters demonstrated their sensitivity to the design.

Linear quadratic optimal model‐following control of a helicopter in hover

Linear quadratic optimal model‐following control of a helicopter in hover

Evaluation of Simulation Motion Fidelity Criteria in the Vertical and Directional Axes

An evaluation of existing motion fidelity criteria was conducted on the NASA Ames Vertical Motion Simulator. Experienced test pilots flew single-axis repositioning tasks in both the vertical and the directional axes. Using a first-order approximation of a hovering helicopter, tasks were flown with variations only in the filters that attenuate the commands to the simulator motion system. These filters had second-order high-pass characteristics, and the variations were made in the filter gain and natural frequency. The variations spanned motion response characteristics from nearly full math-model motion to fixed-base. Between configurations, pilots recalibrated their motion response perception by flying the task with full motion. Pilots subjectively rated the motion fidelity of subsequent configurations relative to this full motion case, which was considered the standard for comparison. The results suggested that the existing vertical-axis criterion was accurate for combinations of gain and natural frequency changes. However, if only the gain or the natural frequency was changed, the rated motion fidelity was better than the criterion predicted. In the vertical axis, the objective and subjective results indicated that a larger gain reduction was tolerated than the existing criterion allowed. The limited data collected in the yaw axis revealed that pilots had difficulty in distinguishing among the variations in the pure yaw motion cues.

Vortex and momentum theories for hovering rotors

ILLER1 has called attention to a seeming paradox in the vortex theory for hovering rotors. In the usual application of vortex theory for the limiting case of an infinite number of rotor blades to propellers, the induced velocities in the wake are small compared to the forward velocity and the distributed vortices in the wake may be assumed to move downstream at the propeller forward velocity for lightly loaded propellers. Theodorsen showed,2 however, that in the application of the Goldstein theory for optimum circulation distribution on propellers with a finite number of blades, it is necessary, for more heavily loaded propellers representative of practical designs, to take into account the induced velocities in arriving at vortex geometry in the wake. Theodorsen also gave an approximate method for computing slipstream contraction, but for the cases of propellers in cruise flight which he considered, the slipstream contraction was of the order of 1%; the analysis specifically excluded static thrust conditions. For propeller cruise operating conditions it is well known3'4 that the results of vortex theory for an infinite number of blades are essentially identical to the momentum theory of the actuator disk; this remains true for the case of nonconstant blade circulation if both theories are applied to differential annular strips on the blades. It is also true when induced wake rotation is introduced into both theories. A particular result of practical significance in the theories (neglecting wake rotation) is that the induced axial velocity at the propeller vl is one-half the velocity in the ultimate wake, v2. In the propeller case, with negligible slipstream contraction, this follows from the fact that the vortices are assumed to be uniformly distributed over the circumference of a cylinder of approximately constant diameter that has a semi-infinite length viewed from the plane of the propeller and an infinite length viewed from a transverse plane in the ultimate wake, giving rise to a factor of two in the calculated induced velocities. It should be noted, however, that in the momentum theory of the actuator disk it is assumed that the flow velocity in the wake is constant for constant disk loading. In the vortex theory, this is a conclusion arrived at from analyses of the induced flowfield in the vortex cylinder. (Yet another approach is to represent the acutator disk by a doublet layer; this gives results equivalent to vortex theory.) With large slipstream contractions, such as occur for the static thrust condition of a propeller or for a hovering helicopter rotor, it is no longer valid to make this particular calculation since the wake vortices cannot be considered to lie on cylinders of constant diameter in the vicinity of the propeller plane. Yet, even in this case, vortex theory for an

A Model of the Unsteady Aerodynamics of a Hovering Helicopter Rotor That Includes Variations of the Wake Geometry

A Model of the Unsteady Aerodynamics of a Hovering Helicopter Rotor That Includes Variations of the Wake Geometry

A New Model of Rotor Dynamics During Pitch and Roll of a Hovering Helicopter

A New Model of Rotor Dynamics During Pitch and Roll of a Hovering Helicopter

Identification and Simulation Evaluation of a Combat Helicopter in Hover

Frequency-domain parameter identification techniques were used to develop a hover mathematical model of the AH-64 Apache helicopter from flight data. The unstable AH-64 bare-airframe characteristics, without a stability augmentation system, were parameterized in the conventional stability-derivative form. To improve the model's vertical response, a simple transfer-function model approximating the effects of dynamic inflow was developed. The model, with and without stability augmentation, was then evaluated by AH-64 pilots in a moving-base simulation. It was the opinion of the pilots that the simulation was a satisfactory representation of the aircraft for the tasks of interest. The principal negative comment was that height control was more difficult in the simulation than in the aircraft.

Rotor‐State Feedback in the Design of Flight Control Laws for a Hovering Helicopter

The use of rigid-body and rotor-state feedback gains in the design of helicopter flight control laws was investigated analytically on a blade element, articulated rotor, helicopter model. The study was conducted while designing a control law to meet an existing military rotorcraft handling qualities design specification (ADS-33C) in low-speed flight. A systematic approach to meet this specification was developed along with an assessment of the function of these gains in the feedback loops. Using the results of this assessment, the pitch and roll crossover behavior was easily modified by adjusting the body attitude and rotor-flap feedback gains. Critical to understanding the feedback gains is that the roll and pitch rate dynamics each have second-order behavior, not the classic first-order behavior, which arises from a quasi-static rotor, six degree-of-freedom model.

Synthesis and evaluation of an H2 control law for a hovering helicopter

The design and simulator evaluation of a rate-command flight-control law for a UH-60 helicopter in near-hover flight conditions are described. The multivariable control law was synthesized using an //2 method, which, through weighting functions, directly shapes the singular values of the sensitivity (/ + GK)~l, comple- mentary sensitivity GK(I + GK)~ l, and control input K(I + GK)~l transfer-function matrices. The design was implemented on the vertical motion simulator, and four low-speed hover tasks were used to evaluate the control system characteristics. The pilot comments indicated good decoupling and quick response characteristics, but also revealed a mild pilot-induced oscillation tendency in the roll axis. HE purpose of the research documented in this paper is to investigate the application of an H2 multivariable de- sign method to synthesize flight-control laws for helicopters. The objective of the work is to determine through analysis and piloted simulation the strengths and weaknesses of the H2 method within the context of a modern rotorcraft design specification. The choice of a multivariable design method is natural to the problem presented because of the highly coupled nature of helicopter dynamics and the desired level of feedback dictated by modern design specifications. Helicopter flight-control laws designed through classical single-input, single-output techniques are satisfactory when requirements do not infringe on the frequency range of significant interaxis cross coupling due to the rotor. Previous generation helicopters and their design requirements have complied with this assumption. However, future helicopters will use hingeless and bearingless rotors with more complex coupling, and these helicopters will have control-law design requirements demanding higher levels of feedback. To ensure that the design criteria are sufficiently stringent to represent future requirements, the control law here was de- signed according to the Section 3 hover requirements in the Aeronautical Design Standard (ADS-33C).1 Previous work on multivariable techniques for rotorcraft flight-control design2'4 complied with part or none of the ADS-33C requirements, making these works difficult to interpret in terms of this new standard. These previous works also used a low-order model (usually a 6-degree-of-freedom model) to design the flight-control law, and the final design was usually checked on a higher-order model in analysis or simulation. These higher-order models did not fully represent the rotor-fuselage dynamics, lacking either the coupled nature of the rotor dynamics (between the fuselage and the rotor and/or between the rotor degrees of freedom) or lacking the rotor lead-lag degree of freedom. These effects are known limiting factors in the feedback gains of helicopter flight-control laws and are represented in the analysis and piloted simulator model here. The importance of these effects was illustrated in advanced digital optical control system (ADOCS) research, where feedback gain reductions of 40% were necessary to avoid lead-lag instabilities.5 This paper continues with a description of the models used in the design process and piloted simulation, followed by a description of the method used to synthesize the flight-control law. The salient features of a rate-command control-law de- sign are then discussed, followed by descriptions of the simu- lation environment and task arrangement. Finally, the results of the simulation are discussed and conclusions drawn. More details on the design method, design models, and flight-con- trol design can be found in Ref. 6, whereas more details related to the simulation experiment can be located in Ref. 7.

Tip Vortex Geometry of a Hovering Helicopter Rotor in Ground Effect

The wide-field shadowgraph method has been used to photograph the tip vortices of a hovering helicopter rotor in ground effect. The shadowgraphs were used to obtain quantitative measurements of the rotor tip vortex geometry both in and out of ground effect. Many important phenomena are visible in the rotor wake using this method. These include the variation in descent and contraction rates of the tip vortices in ground effect, and the interaction between tip vortices in the far wake. The tip vortex geometry from the shadowgraphs is compared with the tip vortex geometry predicted using a free wake hover performance analysis. The free wake analysis accurately predicts the tip vortex geometry both in and out of ground effect. Performance data from the test is compared with the performance predicted using several methods, including the free wake analysis. All methods provided reasonable predictions for the helicopter performance in ground effect.

A Control Project for Electronics Engineering Students

A control project for electronics engineering students The relevance of control theory to electrical engineering can be demonstrated vividly to undergraduate students by its application to the design of linear continuous electronic circuits that control the height of a hovering helicopter, animated on a PC screen. The object of the student project is to design altitude control electronics by using control theory. The flexibility of the PC allows for full data logging and graphic display features as well as giving each student a unique set of parameters.

Applying the lifting surface method to the blade tip of a hovering helicopter

The approach of inner/outer fields is applied in order to calculate the aerodynamic loads at a blade tip (inner field) of a helicopter in hovering or axial flight. It is shown that if the inner field includes the wake in the vicinity of the tip region, then the well known lifting surface method for fixed wings can be applied in order to study the aerodynamic field at the tip. This method is more efficient than other methods that have been previously applied and thus it is suitable for repeated analyses during the design procedure. Furthermore, the method is capable of dealing with complicated tip geometries without requiring excessive computer resources. Examples of using the method are presented, where the calculated results are compared with experimental ones.

The Evaluation of a Feral Pig Eradication Program During a Simulated Exotic Disease Outbreak.

An exercise was conducted to evaluate the effectiveness of plans to eradicate feral pigs in an exotic disease emergency. The study site was an area of 120 km2 on the southern edge of the Macquarie Marshes in western New South Wales. Shooting from a helicopter accounted for 946 pigs at a rate of 39.2 per hour. This was at an average of 1.65 shots and a cost of $11.77 per pig. A further 43 were shot from the ground or trapped. Of an estimated initial population of 1238, 80% was removed. Telemetry studies conducted in conjunction with the exercise indicated that some pigs became attuned to the significance of a hovering helicopter and modified their behaviour to avoid detection. Movements also emphasised the need to match the boundaries of feral pig eradication zones with natural boundaries, where overlapping home ranges are minimal and densities low. Eradication of feral pigs during an outbreak of exotic disease may be an unrealistic goal, and it may be more efficient to aim to eradicate the disease within the feral pig population. This would be achieved by isolating those pigs carrying the infection; it does not necessarily require the removal of all feral pigs.

A Simplified Free Wake Method for Horizontal-Axis Wind Turbine Performance Prediction

Based on the assumption that wake geometry of a horizontal-axis wind turbine closely resembles that of a hovering helicopter, a method is presented for predicting the performance of a horizontal-axis wind turbine. A vortex method is used in which the wake is composed of an intense tip-vortex and a diffused inboard wake. Performance parameters are calculated by application of the Biot-Savart law along with the Kutta-Joukowski theorem. Predictions are shown to compare favorably with values from a more complicated full free wake analysis and with existing experimental data, but require more computational effort than an existing fast free wake method.