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  • Inorganic Chemistry  (7,167)
  • Aircraft Stability and Control
  • 1995-1999  (1,992)
  • 1975-1979  (3,765)
  • 1950-1954  (1,662)
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  • 1
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    Closed-Loop System Identification Experience for Flight Control Law and Flying Qualities Evaluation of a High Performance Fighter Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2011-09-23
    Description: This paper highlights some of the results and issues associated with estimating models to evaluate control law design methods and design criteria for advanced high performance aircraft. Experimental fighter aircraft such as the NASA High Alpha Research Vehicle (HARV) have the capability to maneuver at very high angles of attack where nonlinear aerodynamics often predominate. HARV is an experimental F/A-18, configured with thrust vectoring and conformal actuated nose strakes. Identifying closed-loop models for this type of aircraft can be made difficult by nonlinearities and high-order characteristics of the system. In this paper only lateral-directional axes are considered since the lateral-directional control law was specifically designed to produce classical airplane responses normally expected with low-order, rigid-body systems. Evaluation of the control design methodology was made using low-order equivalent systems determined from flight and simulation. This allowed comparison of the closed-loop rigid-body dynamics achieved in flight with that designed in simulation. In flight, the On Board Excitation System was used to apply optimal inputs to lateral stick and pedals at five angles of attack: 5, 20, 30, 45, and 60 degrees. Data analysis and closed-loop model identification were done using frequency domain maximum likelihood. The structure of the identified models was a linear state-space model reflecting classical 4th-order airplane dynamics. Input time delays associated with the high-order controller and aircraft system were accounted for in data preprocessing. A comparison of flight estimated models with small perturbation linear design models highlighted nonlinearities in the system and indicated that the estimated closed-loop rigid-body dynamics were sensitive to input amplitudes at 20 and 30 degrees angle of attack.
    Keywords: Aircraft Stability and Control
    Type: System Identification for Integrated Aircraft Development and Flight Testing; 16-1 - 16-13; RTO-MP-11
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  • 2
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    High Reynolds Number Effects on HSCT Stability and Control Characteristics (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Two wind tunnel tests during 1995 in the National Transonic Facility (NTF 070 and 073) served to define Reynolds number effects on longitudinal and lateral-directional stability and control. Testing was completed at both high lift and transonic conditions. The effect of Reynolds number on the total airplane configuration, horizontal and vertical tail effectiveness, forebody chine performance, rudder control and model aeroelastics was investigated. This paper will present pertinent stability and control results from these two test entries. Note that while model aeroelastic effects are examined in this presentation, no corrections for these effects have been made to the data.
    Keywords: Aircraft Stability and Control
    Type: First NASA/Industry High-Speed Research Configuration Aerodynamics Workshop; Part 3; 1253-1284; NASA/CP-1999-209690/PT3
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  • 3
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    Correlation of Fin Buffet Pressures on an F/A-18 with Scaled Wind-Tunnel Measurements (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Buffeting is an aeroelastic phenomenon occurring at high angles of attack that plagues high performance aircraft, especially those with twin vertical tails. Previous wind-tunnel and flight tests were conducted to characterize the buffet loads on the vertical tails by measuring surface pressures, bending moments, and accelerations. Following these tests, buffeting responses were computed using the measured buffet pressures and compared to the measured buffeting responses. The calculated results did not match the measured data because the assumed spatial correlation of the buffet pressures was not correct. A better understanding of the partial (spatial) correlation of the differential buffet pressures on the tail was necessary to improve the buffeting predictions. Several wind-tunnel investigations were conducted for this purpose. When compared, the results of these tests show that the partial correlation scales with flight conditions. One of the remaining questions is whether the wind-tunnel data is consistent with flight data. Presented herein, cross-spectra and coherence functions calculated from pressures that were measured on the High Alpha Research Vehicle indicate that the partial correlation of the buffet pressures in flight agrees with the partial correlation observed in the wind tunnel.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 2; 615-626; NASA/CP-1999-209136/PT2
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  • 4
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    An Experimental Database for Conventional and Alternate Control Concepts on the HSR 1.675% Reference H Model (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The objective was to experimentally evaluate the longitudinal and lateral-directional stability and control characteristics of the Reference H configuration at supersonic and transonic speeds. A series of conventional and alternate control devices were also evaluated at supersonic and transonic speeds. A database on the conventional and alternate control devices was to be created for use in the HSR program.
    Keywords: Aircraft Stability and Control
    Type: First NASA/Industry High-Speed Research Configuration Aerodynamics Workshop; Part 3; 1233-1251; NASA/CP-1999-209690/PT3
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  • 5
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    Critical Stability and Control Issues in High-Speed Aerodynamics for the HSCT (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The stability and control issues in high speed aerodynamics of most significance for the development of a viable HSCT are identified, and the status of the Ref. H configuration with respect to these issues is discussed. The interdependence between aerodynamic requirements and assumptions about airplane system functions such as Envelope Protection and Integrated Flight/Propulsion Control is highlighted. The conclusions presented draw on results from the Ref. H Assessment and Alternate Control Concepts Assessment performed under Configuration Aerodynamics Subtask 5 during 1995.
    Keywords: Aircraft Stability and Control
    Type: First NASA/Industry High-Speed Research Configuration Aerodynamics Workshop; Part 3; 1215-1231; NASA/CP-1999-209690/PT3
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  • 6
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    An Examination of Applying Shunted Piezoelectrics to Reduce Aeroelastic Response (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Several analytical and experimental studies clearly demonstrate that piezoelectric materials (piezoelectrics) can be used as actuators to actively control vibratory response, including aeroelastic response. However, two important issues in using piezoelectrics as actuators for active control are: 1) the potentially large amount of power required to operate the actuators, and 2) the complexities involved with active control (added hardware, control law design, and implementation). Active or passive damping augmentation using shunted piezoelectrics may provide a viable alternative. This approach requires only simple electrical circuitry and very little or no electrical power. The current study examines the feasibility of using shunted piezoelectrics to reduce aeroelastic response using a typical-section representation of a wing and piezoelectrics shunted with a parallel resistor and inductor. The aeroelastic analysis shows that shunted piezoelectrics can effectively reduce aeroelastic response below flutter and may provide a simple, low-power method of subcritical aeroelastic control.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 2; 553-572; NASA/CP-1999-209136/PT2
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  • 7
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    An H-infinity Approach to Control Synthesis with Load Minimization for the F/A-18 Active Aeroelastic Wing (1999)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H(sub infinity) synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.
    Keywords: Aircraft Stability and Control
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  • 8
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    Aeroelastic Tailoring for Stability Augmentation and Performance Enhancements of Tiltrotor Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The requirements for increased speed and productivity for tiltrotors has spawned several investigations associated with proprotor aeroelastic stability augmentation and aerodynamic performance enhancements. Included among these investigations is a focus on passive aeroelastic tailoring concepts which exploit the anisotropic capabilities of fiber composite materials. Researchers at Langley Research Center and Bell Helicopter have devoted considerable effort to assess the potential for using these materials to obtain aeroelastic responses which are beneficial to the important stability and performance considerations of tiltrotors. Both experimental and analytical studies have been completed to examine aeroelastic tailoring concepts for the tiltrotor, applied either to the wing or to the rotor blades. This paper reviews some of the results obtained in these aeroelastic tailoring investigations and discusses the relative merits associated with these approaches.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 1; 121-138; NASA/CP-1999-209136/PT1
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  • 9
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    An H-Infinity Approach to Control Synthesis with Load Minimization for the F/A-18 Active Aeroelastic Wing (1999)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H-infinity synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 1; 23-32; NASA/CP-1999-209136/PT1
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  • 10
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    Wavelet Applications for Flight Flutter Testing (1999)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Wavelets present a method for signal processing that may be useful for analyzing responses of dynamical systems. This paper describes several wavelet-based tools that have been developed to improve the efficiency of flight flutter testing. One of the tools uses correlation filtering to identify properties of several modes throughout a flight test for envelope expansion. Another tool uses features in time-frequency representations of responses to characterize nonlinearities in the system dynamics. A third tool uses modulus and phase information from a wavelet transform to estimate modal parameters that can be used to update a linear model and reduce conservatism in robust stability margins.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 1; 393-402; NASA/CP-1999-209136/PT1
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  • 11
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    Transonic Flutter Suppression Control Law Design, Analysis and Wind Tunnel Results (1999)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 1; 381-392; NASA/CP-1999-209136/PT1
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  • 12
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    Neural Networks Used to Compare Designed and Measured Time-Average Patterns (1999)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Electronic time-average holograms are convenient for comparing the measured vibration modes of fan blades with those calculated by finite-element models. At the NASA Lewis Research Center, neural networks recently were trained to perform what had been a simple visual comparison of the predictions of the design models with the measurements. Finite-element models were used to train neural networks to recognize damage and strain information encoded in subtle changes in the time-average patterns of cantilevers. But the design-grade finite element models were unable to train the neural networks to detect damage in complex blade shapes. The design-model-generated patterns simply did not agree well enough with the measured patterns. Instead, hybrid-training records, with measured time-average patterns as the input and model-generated strain information as the output, were used to effect successful training.
    Keywords: Aircraft Stability and Control
    Type: Research and Technology 1998; NASA/TM-1999-208815
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  • 13
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    Experimental Investigation of Convoluted Contouring for Aircraft Afterbody Drag Reduction (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An experimental investigation was performed in the NASA Langley 16-Foot Transonic Tunnel to determine the aerodynamic effects of external convolutions, placed on the boattail of a nonaxisymmetric nozzle for drag reduction. Boattail angles of 15 and 22 were tested with convolutions placed at a forward location upstream of the boattail curvature, at a mid location along the curvature and at a full location that spanned the entire boattail flap. Each of the baseline nozzle afterbodies (no convolutions) had a parabolic, converging contour with a parabolically decreasing corner radius. Data were obtained at several Mach numbers from static conditions to 1.2 for a range of nozzle pressure ratios and angles of attack. An oil paint flow visualization technique was used to qualitatively assess the effect of the convolutions. Results indicate that afterbody drag reduction by convoluted contouring is convolution location, Mach number, boattail angle, and NPR dependent. The forward convolution location was the most effective contouring geometry for drag reduction on the 22 afterbody, but was only effective for M 〈 0.95. At M = 0.8, drag was reduced 20 and 36 percent at NPRs of 5.4 and 7, respectively, but drag was increased 10 percent for M = 0.95 at NPR = 7. Convoluted contouring along the 15 boattail angle afterbody was not effective at reducing drag because the flow was minimally separated from the baseline afterbody, unlike the massive separation along the 22 boattail angle baseline afterbody.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-2670 , 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jun 20, 1999 - Jun 24, 1999; Los Angeles, CA; United States
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  • 14
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    Transonic Flutter Suppression Control Law Design Using Classical and Optimal Techniques with Wind-Tunnel Results (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-1396 , 40th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials (SDM) Conference; Apr 12, 1999 - Apr 15, 1999; Saint Louis, MO; United States
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  • 15
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    Numerical Modeling Studies of Wake Vortices: Real Case Simulations (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A three-dimensional large-eddy simulation model, TASS, is used to simulate the behavior of aircraft wake vortices in a real atmosphere. The purpose for this study is to validate the use of TASS for simulating the decay and transport of wake vortices. Three simulations are performed and the results are compared with the observed data from the 1994-1995 Memphis field experiments. The selected cases have an atmospheric environment of weak turbulence and stable stratification. The model simulations are initialized with appropriate meteorological conditions and a post roll-up vortex system. The behavior of wake vortices as they descend within the atmospheric boundary layer and interact with the ground is discussed.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-0755 , 37th AIAA Aerospace Sciences Meeting and Exhibit; Jan 11, 1999 - Jan 14, 1999; Reno, NV; United States
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  • 16
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    Effect of Dynamic Rolling Oscillations on Twin Tail Buffet Response (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The effect of dynamic rolling oscillations of delta-wing/twin-tail configuration on twin-tail buffet response is investigated. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. The configuration model is statically pitched at 30 deg. angle of attack and then forced to oscillate in roll around the symmetry axis at a constant amplitude of 4 deg. and reduced frequency of pi and 2(pi). The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. This multidisciplinary problem is solved using three sets of equations on a dynamic multi-block grid structure. The first set is the unsteady, full Navier-Stokes equations, the second set is the aeroelastic equations for coupled bending and torsion vibrations of the tails, and the third set is the grid-displacement equations. The configuration is investigated for inboard position of the twin tails which corresponds to a separation distance between the twin tails of 33% wing span. The computed results are compared with the results of stationary configuration, which previously have been validated using experimental data. The results conclusively showed that the rolling oscillations of the configuration have led to higher loads, higher deflections, and higher excitation peaks than those of the stationary configuration. Moreover, increasing the reduced frequency has led to higher loads and excitation peaks and lower bending and torsion deflections and acceleration.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-0792 , Aerospace Sciences; Jan 11, 1999 - Jan 14, 1999; Reno, NV; United States
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  • 17
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    X-33 Ascent Flight Controller Design by Trajectory Linearization: A Singular Perturbational Approach (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-08-16
    Description: The flight control of X-33 poses a challenge to conventional gain-scheduled flight controllers due to its large attitude maneuvers from liftoff to orbit and reentry. In addition, a wide range of uncertainties in vehicle handling qualities and disturbances must be accommodated by the attitude control system. Nonlinear tracking and decoupling control by trajectory linearization can be viewed as the ideal gain-scheduling controller designed at every point on the flight trajectory. Therefore it provides robust stability and performance at all stages of flight without interpolation of controller gains, and eliminates costly controller redesigns due to minor airframe alteration or mission reconfiguration. A prototype trajectory linearization design for X-33 ascent flight controller was designed and tested with 3-DOF and 6-DOF simulations during the 10 weeks SFFP. It is noted that the 6-DOF results were obtained from the 3-DOF design with only a few hours of tuning, which demonstrates the inherent robustness of the design technique. It is this "plug-and-play" feature that is much needed by NASA for the development, test and routine operations of the RLVs. Plans for further research are also presented.
    Keywords: Aircraft Stability and Control
    Type: 1999 NASA/ASEE Summer Faculty Fellowship Program; D-53
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  • 18
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    Determination of Stability and Control Derivatives using Computational Fluid Dynamics and Automatic Differentiation (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: With the recent interest in novel control effectors there is a need to determine the stability and control derivatives of new aircraft configurations early in the design process. These derivatives are central to most control law design methods and would allow the determination of closed-loop control performance of the vehicle. Early determination of the static and dynamic behavior of an aircraft may permit significant improvement in configuration weight, cost, stealth, and performance through multidisciplinary design. The classical method of determining static stability and control derivatives - constructing and testing wind tunnel models - is expensive and requires a long lead time for the resultant data. Wind tunnel tests are also limited to the preselected control effectors of the model. To overcome these shortcomings, computational fluid dynamics (CFD) solvers are augmented via automatic differentiation, to directly calculate the stability and control derivatives. The CFD forces and moments are differentiated with respect to angle of attack, angle of sideslip, and aircraft shape parameters to form these derivatives. A subset of static stability and control derivatives of a tailless aircraft concept have been computed by two differentiated inviscid CFD codes and verified for accuracy with central finite-difference approximations and favorable comparisons to a simulation database.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-3136
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  • 19
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    Application of Forebody Strakes for Directional Stability and Control of Transport Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A brief overview of a cooperative NASA/Boeing research effort, Strake Technology Research Application to Transport Aircraft (STRATA), intended to explore the potential of applying forebody strake technology to transport aircraft configurations for directional stability and control at low angles of attack, is presented. As an initial step in the STRATA program, an exploratory wind-tunnel investigation of the effect of fixed forebody strakes on the directional stability and control characteristics of a generic transport configuration was conducted in the NASA Langley 12-Foot Low-Speed Wind Tunnel. Results of parametric variations in strake chord and span, as well as the effect of strake incidence, are presented. The use of strakes for yaw control is also discussed. Results emphasize the importance of forebody/fuselage crossflow in influencing strake effectiveness. Strake effectiveness is also seen to be directly related to its span, but less sensitive to chord; a very short-chord strake with sufficient span can have a significant effect.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-4448
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  • 20
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    Control Law Synthesis for Vertical Fin Buffeting Alleviation Using Strain Actuation (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: In the present investigation, the results obtained during the ground test of a closed-loop control system conducted on a full-scale fighter to attenuate vertical fin buffeting response using strain actuation are presented. Two groups of actuators consisting of piezoelectric elements distributed over the structure were designed to achieve authority over the first and second modes of the vertical fin. The control laws were synthesized using the Linear Quadratic Gaussian (LQG) method for a time-invariant control system. Three different pairs of sensors including strain gauges and accelerometers at different locations were used to close the feedback loop. The results demonstrated that measurable reductions in the root-mean-square (RMS) values of the fin dynamic response identified by the strain transducer at the critical point for fatigue at the root were achieved under the most severe buffet condition. For less severe buffet conditions, reductions of up to 58% were achieved.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-1317
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  • 21
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    Investigation of Dynamic Flight Maneuvers With an Iced Tailplane (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A detailed analysis of two of the dynamic maneuvers, the pushover and elevator doublet, from the NASA/FAA Tailplane Icing Program are discussed. For this series of flight tests, artificial ice shapes were attached to the leading edge of the horizontal stabilizer of the NASA Lewis Research Center icing aircraft, a DHC-6 Twin Otter. The purpose of these tests was to learn more about ice-contaminated tailplane stall (ICTS), the known cause of 16 accidents resulting in 139 fatalities. The pushover has been employed by the FAA, JAA and Transport Canada for tailplane icing certification. This research analyzes the pushover and reports on the maneuver performance degradation due to ice shape severity and flap deflection. A repeatability analysis suggests tolerances for meeting the required targets of the maneuver. A second maneuver, the elevator doublet, is also studied.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-208849 , E-11470 , NAS 1.15:208849 , AIAA Paper 99-0371 , Aerospace Sciences; Jan 11, 1999 - Jan 14, 1999; Reno, NV; United States
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  • 22
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    Reference H Cycle 3 Stability, Control, and Flying Qualities Batch Assessments (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This work is an update of the assessment completed in February of 1996, when a preliminary assessment report was issued for the Cycle 2B simulation model. The primary purpose of the final assessment was to re-evaluate each assessment against the flight control system (FCS) requirements document using the updated model. Only a limited number of final assessments were completed due to the close proximity of the release of the Langley model and the assessment deliverable date. The assessment used the nonlinear Cycle 3 simulation model because it combines nonlinear aeroelastic (quasi-static) aerodynamic with hinge moment and rate limited control surface deflections. Both Configuration Aerodynamics (Task 32) and Flight Controls (Task 36) were funded in 1996 to conduct the final stability and control assessments of the unaugmented Reference H configuration in FY96. Because the two tasks had similar output requirements, the work was divided such that Flight Controls would be responsible for the implementation and checkout of the simulation model and Configuration Aerodynamics for writing Madab "script' files, conducting the batch assessments and writing the assessment report. Additionally, Flight Controls was to investigate control surface allocations schemes different from the baseline Reference H in an effort to fulfill flying qualities criteria.
    Keywords: Aircraft Stability and Control
    Type: 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop; 1; Part 1; 441-476; NASA/CP-1999-209691/VOL1/PT1
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  • 23
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    Development of An Intelligent Flight Propulsion Control System (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The initial design and demonstration of an Intelligent Flight Propulsion and Control System (IFPCS) is documented. The design is based on the implementation of a nonlinear adaptive flight control architecture. This initial design of the IFPCS enhances flight safety by using propulsion sources to provide redundancy in flight control. The IFPCS enhances the conventional gain scheduled approach in significant ways: (1) The IFPCS provides a back up flight control system that results in consistent responses over a wide range of unanticipated failures. (2) The IFPCS is applicable to a variety of aircraft models without redesign and,(3) significantly reduces the laborious research and design necessary in a gain scheduled approach. The control augmentation is detailed within an approximate Input-Output Linearization setting. The availability of propulsion only provides two control inputs, symmetric and differential thrust. Earlier Propulsion Control Augmentation (PCA) work performed by NASA provided for a trajectory controller with pilot command input of glidepath and heading. This work is aimed at demonstrating the flexibility of the IFPCS in providing consistency in flying qualities under a variety of failure scenarios. This report documents the initial design phase where propulsion only is used. Results confirm that the engine dynamics and associated hard nonlineaaities result in poor handling qualities at best. However, as demonstrated in simulation, the IFPCS is capable of results similar to the gain scheduled designs of the NASA PCA work. The IFPCS design uses crude estimates of aircraft behaviour. The adaptive control architecture demonstrates robust stability and provides robust performance. In this work, robust stability means that all states, errors, and adaptive parameters remain bounded under a wide class of uncertainties and input and output disturbances. Robust performance is measured in the quality of the tracking. The results demonstrate the flexibility of the IFPCS architecture and the ability to provide robust performance under a broad range of uncertainty. Robust stability is proved using Lyapunov like analysis. Future development of the IFPCS will include integration of conventional control surfaces with the use of propulsion augmentation, and utilization of available lift and drag devices, to demonstrate adaptive control capability under a greater variety of failure scenarios. Further work will specifically address the effects of actuator saturation.
    Keywords: Aircraft Stability and Control
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    Reconfigurable Flight Control Designs With Application to the X-33 Vehicle (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Two methods for control system reconfiguration have been investigated. The first method is a robust servomechanism control approach (optimal tracking problem) that is a generalization of the classical proportional-plus-integral control to multiple input-multiple output systems. The second method is a control-allocation approach based on a quadratic programming formulation. A globally convergent fixed-point iteration algorithm has been developed to make onboard implementation of this method feasible. These methods have been applied to reconfigurable entry flight control design for the X-33 vehicle. Examples presented demonstrate simultaneous tracking of angle-of-attack and roll angle commands during failures of the right body flap actuator. Although simulations demonstrate success of the first method in most cases, the control-allocation method appears to provide uniformly better performance in all cases.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-206582 , H-2345 , NAS 1.15:206582 , AIAA Paper 99-4134 , Guidance Navigation and Control; Aug 09, 1999 - Aug 11, 1999; Portland, OR; United States
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    Advanced Aerodynamic Control Effectors (1999)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A 1990 research program that focused on the development of advanced aerodynamic control effectors (AACE) for military aircraft has been reviewed and summarized. Data are presented for advanced planform, flow control, and surface contouring technologies. The data show significant increases in lift, reductions in drag, and increased control power, compared to typical aerodynamic designs. The results presented also highlighted the importance of planform selection in the design of a control effector suite. Planform data showed that dramatic increases in lift (greater than 25%) can be achieved with multiple wings and a sawtooth forebody. Passive porosity and micro drag generator control effector data showed control power levels exceeding that available from typical effectors (moving surfaces). Application of an advanced planform to a tailless concept showed benefits of similar magnitude as those observed in the generic studies.
    Keywords: Aircraft Stability and Control
    Type: SAE-1999-01-5619 , 1999 World Aviation Congress; Oct 19, 1999 - Oct 21, 1999; San Francisco, CA; United States
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    Wavelet Applications for Flight Flutter Testing (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wavelets present a method for signal processing that may be useful for analyzing responses of dynamical systems. This paper describes several wavelet-based tools that have been developed to improve the efficiency of flight flutter testing. One of the tools uses correlation filtering to identify properties of several modes throughout a flight test for envelope expansion. Another tool uses features in time-frequency representations of responses to characterize nonlinearities in the system dynamics. A third tool uses modulus and phase information from a wavelet transform to estimate modal parameters that can be used to update a linear model and reduce conservatism in robust stability margins.
    Keywords: Aircraft Stability and Control
    Type: H-2364 , International Forum on Aeroelasticity and Structural Dynamics; Jun 22, 1999 - Jun 25, 1999; Williamsburg, VA; United States
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    Initial Flight Test of the Production Support Flight Control Computers at NASA Dryden Flight Research Center (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The NASA Dryden Flight Research Center has completed the initial flight test of a modified set of F/A-18 flight control computers that gives the aircraft a research control law capability. The production support flight control computers (PSFCC) provide an increased capability for flight research in the control law, handling qualities, and flight systems areas. The PSFCC feature a research flight control processor that is "piggybacked" onto the baseline F/A-18 flight control system. This research processor allows for pilot selection of research control law operation in flight. To validate flight operation, a replication of a standard F/A-18 control law was programmed into the research processor and flight-tested over a limited envelope. This paper provides a brief description of the system, summarizes the initial flight test of the PSFCC, and describes future experiments for the PSFCC.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-206581 , H-2343 , NAS 1.15:206581 , AIAA Paper 99-4203 , Guidance, Navigation, and Control; Aug 09, 1999 - Aug 11, 1999; Portland, OR; United States
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    Transonic Flutter Suppression Control Law Design, Analysis and Wind-Tunnel Results (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.
    Keywords: Aircraft Stability and Control
    Type: IFA-1999 , Aeroelasticity and Structural Dynamics 1999; Jun 22, 1999 - Jun 25, 1999; Williamsburg, VA; United States
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    Merged Vision and GPS Control of a Semi-Autonomous, Small Helicopter (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This final report documents the activities performed during the research period from April 1, 1996 to September 30, 1997. It contains three papers: Carrier Phase GPS and Computer Vision for Control of an Autonomous Helicopter; A Contestant in the 1997 International Aerospace Robotics Laboratory Stanford University; and Combined CDGPS and Vision-Based Control of a Small Autonomous Helicopter.
    Keywords: Aircraft Stability and Control
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  • 30
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    Transonic Flutter Suppression Control Law Design Using Classical and Optimal Techniques with Wind-Tunnel Results (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-1396 , Structures, Structural Dynamics and Materials; Apr 12, 1999 - Apr 15, 1999; Saint Louis, MO; United States
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    Flight Test of an Adaptive Configuration Optimization System for Transport Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A NASA Dryden Flight Research Center program explores the practical application of real-time adaptive configuration optimization for enhanced transport performance on an L-1011 aircraft. This approach is based on calculation of incremental drag from forced-response, symmetric, outboard aileron maneuvers. In real-time operation, the symmetric outboard aileron deflection is directly optimized, and the horizontal stabilator and angle of attack are indirectly optimized. A flight experiment has been conducted from an onboard research engineering test station, and flight research results are presented herein. The optimization system has demonstrated the capability of determining the minimum drag configuration of the aircraft in real time. The drag-minimization algorithm is capable of identifying drag to approximately a one-drag-count level. Optimizing the symmetric outboard aileron position realizes a drag reduction of 2-3 drag counts (approximately 1 percent). Algorithm analysis of maneuvers indicate that two-sided raised-cosine maneuvers improve definition of the symmetric outboard aileron drag effect, thereby improving analysis results and consistency. Ramp maneuvers provide a more even distribution of data collection as a function of excitation deflection than raised-cosine maneuvers provide. A commercial operational system would require airdata calculations and normal output of current inertial navigation systems; engine pressure ratio measurements would be optional.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-206569 , NAS 1.15:206569 , H-2284 , AIAA Paper 99-0831 , Aerospace Sciences; Jan 11, 1999 - Jan 14, 1999; Reno, NV; United States
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    User's Guide for ENSAERO_FE Parallel Finite Element Solver (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A high fidelity parallel static structural analysis capability is created and interfaced to the multidisciplinary analysis package ENSAERO-MPI of Ames Research Center. This new module replaces ENSAERO's lower fidelity simple finite element and modal modules. Full aircraft structures may be more accurately modeled using the new finite element capability. Parallel computation is performed by breaking the full structure into multiple substructures. This approach is conceptually similar to ENSAERO's multizonal fluid analysis capability. The new substructure code is used to solve the structural finite element equations for each substructure in parallel. NASTRANKOSMIC is utilized as a front end for this code. Its full library of elements can be used to create an accurate and realistic aircraft model. It is used to create the stiffness matrices for each substructure. The new parallel code then uses an iterative preconditioned conjugate gradient method to solve the global structural equations for the substructure boundary nodes.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-208781 , A-99V0021
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    Reference H Piloted Assessment (LaRC.1) Pilot Briefing Guide (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This document describes the purpose of and method by which an assessment of the Boeing Reference H High-Speed Civil Transport design was evaluated in the NASA Langley Research Center's Visual/Motion Simulator in January 1997. Six pilots were invited to perform approximately 60 different Mission Task Elements that represent most normal and emergency flight operations of concern to the High Speed Research program. The Reference H design represents a candidate configuration for a High-Speed Civil Transport, a second generation supersonic civilian transport aircraft. The High-Speed Civil Transport is intended to be economically sound and environmentally safe while carrying passengers and cargo at supersonic speeds with a trans-Pacific range. This simulation study was designated "LaRC. 1" for the purposes of planning, scheduling and reporting within the Guidance and Flight Controls super-element of the High-Speed Research program. The study was based upon Cycle 3 release of the Reference H simulation model.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-209533 , L-17903 , NAS 1.15:209533
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    Piloted Simulation Investigation of a Supersonic Transport Configuration (LaRC.4) (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This report contains a description of the test facilities and software utilized during a joint NASA/aerospace industry study of improved control laws and desired inceptor characteristics for a candidate supersonic transport air-craft design. Details concerning the characteristics of the simulation cockpit, image generator and display systems, and motion platform are described. Depictions of the various display formats are included. The test schedule, session log, and flight cards describing the maneuvers performed is included. A brief summary of high-lights of the study is given. Modifications made to the industry-provided simulation model are described. This report is intended to serve as a reference document for industry researchers.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-209557 , NAS 1.15:209557
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    Further Development, Support and Enhancement of CONDUIT (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: From the first airplanes steered by handles, wheels, and pedals to today's advanced aircraft, there has been a century of revolutionary inventions, all of them contributing to flight quality. The stability and controllability of aircraft as they appear to a pilot are called flying or handling qualities. Many years after the first airplanes flew, flying qualities were identified and ranked from desirable to unsatisfactory. Later on engineers developed design methods to satisfy these practical criteria. CONDUIT, which stands for Control Designer's Unified Interface, is a modern software package that provides a methodology for optimization of flight control systems in order to improve the flying qualities. CONDUIT is dependent on an the optimization engine called CONSOL-OPTCAD (C-O). C-O performs multicriterion parametric optimization. C-O was successfully tested on a variety of control problems. The optimization-based computational system, C-O, requires a particular control system description as a MATLAB file and possesses the ability to modify the vector of design parameters in an attempt to satisfy performance objectives and constraints specified by the designer, in a C-type file. After the first optimization attempts on the UH-60A control system, an early interface system, named GIFCORCODE (Graphical Interface for CONSOL-OPTCAD for Rotorcraft Controller Design) was created.
    Keywords: Aircraft Stability and Control
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    Gain-Scheduled Approximations to H-Infinity Controllers for the F/A-18 Active Aeroelastic Wing (1999)
    In:  Other Sources
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    Publication Date: 2019-07-10
    Description: The Active Aeroelastic Wing will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This paper presents a control design based on H-infinity synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. The controller is realized as a filter and gain set approximation to a state-space H-infinity controller. This approximation allows scheduling of the controller over a flight envelope.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-4205
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    Adaptive Suction and Blowing for Twin-Tail Buffet Control (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Adaptive active flow control for twin-tail buffet alleviation is investigated. The concept behind this technique is to place control ports on the tail outer and inner surfaces with flow suction or blowing applied through these ports in order to minimize the pressure difference across the tail. The suction or blowing volume flow rate from each port is proportional to the pressure difference across the tail at this location. A parametric study of the effects of the number and location of these ports on the buffet response is carried out. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. This complex multidisciplinary problem is solved sequentially using three sets of equations for the fluid flow, aeroelastic response and grid deformation, using a dynamic multi-block grid structure. The computational model is pitched at 30 deg angle of attack. The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. The model is investigated for the inboard position of the twin tails, which corresponds to a separation distance between the twin tails of 33% of the wing span. Comparison of the time history and power spectral density responses of the tails for various distributions of the control ports are presented and discussed.
    Keywords: Aircraft Stability and Control
    Type: CEAS/AIAA/ICASE/NASA Langley International Forum on Aeroelasticity and Structural Dynamics 1999; Pt. 2; 639-648; NASA/CP-1999-209136/PT2
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    Simulator Evaluation of Simplified Propulsion-Only Emergency Flight Control Systems on Transport Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: With the advent of digital engine control systems, considering the use of engine thrust for emergency flight control has become feasible. Many incidents have occurred in which engine thrust supplemented or replaced normal aircraft flight controls. In most of these cases, a crash has resulted, and more than 1100 lives have been lost. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control capability. Using this PCA system, an F-15 and an MD-11 airplane have been landed without using any flight controls. In simulations, C-17, B-757, and B-747 PCA systems have also been evaluated successfully. These tests used full-authority digital electronic control systems on the engines. Developing simpler PCA systems that can operate without full-authority engine control, thus allowing PCA technology to be installed on less capable airplanes or at lower cost, is also a desire. Studies have examined simplified ?PCA Ultralite? concepts in which thrust control is provided using an autothrottle system supplemented by manual differential throttle control. Some of these concepts have worked well. The PCA Ultralite study results are presented for simulation tests of MD-11, B-757, C-17, and B-747 aircraft.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-206578 , NAS 1.15:206578 , H-2331
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    High Speed Civil Transport Aircraft Simulation: Reference-H Cycle 1, MATLAB Implementation (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The mathematical model and associated code to simulate a high speed civil transport aircraft - the Boeing Reference H configuration - are described. The simulation was constructed in support of advanced control law research. In addition to providing time histories of the dynamic response, the code includes the capabilities for calculating trim solutions and for generating linear models. The simulation relies on the nonlinear, six-degree-of-freedom equations which govern the motion of a rigid aircraft in atmospheric flight. The 1962 Standard Atmosphere Tables are used along with a turbulence model to simulate the Earth atmosphere. The aircraft model has three parts - an aerodynamic model, an engine model, and a mass model. These models use the data from the Boeing Reference H cycle 1 simulation data base. Models for the actuator dynamics, landing gear, and flight control system are not included in this aircraft model. Dynamic responses generated by the nonlinear simulation are presented and compared with results generated from alternate simulations at Boeing Commercial Aircraft Company and NASA Langley Research Center. Also, dynamic responses generated using linear models are presented and compared with dynamic responses generated using the nonlinear simulation.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-209530 , NAS 1.15:209530 , E-17900
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    Cg/Stability Map for the Reference H Cycle 3 Supersonic Transport Concept Along the High Speed Research Baseline Mission Profile (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A comparison is made between the results of trimming a High Speed Civil Transport (HSCT) concept along a reference mission profile using two trim modes. One mode uses the stabilator. The other mode uses fore and aft placement of the center of gravity. A comparison is make of the throttle settings (cruise segments) or the total acceleration (ascent and descent segments) and of the drag coefficient. The comparative stability of trimming using the two modes is also assessed by comparing the stability margins and the placement of the lateral and longitudinal eigenvalues.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209527 , NAS 1.26:209527
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    Multivariable Techniques for High-Speed Research Flight Control Systems (1999)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: This report describes the activities and findings conducted under contract with NASA Langley Research Center. Subject matter is the investigation of suitable multivariable flight control design methodologies and solutions for large, flexible high-speed vehicles. Specifically, methodologies are to address the inner control loops used for stabilization and augmentation of a highly coupled airframe system possibly involving rigid-body motion, structural vibrations, unsteady aerodynamics, and actuator dynamics. Design and analysis techniques considered in this body of work are both conventional-based and contemporary-based, and the vehicle of interest is the High-Speed Civil Transport (HSCT). Major findings include: (1) control architectures based on aft tail only are not well suited for highly flexible, high-speed vehicles, (2) theoretical underpinnings of the Wykes structural mode control logic is based on several assumptions concerning vehicle dynamic characteristics, and if not satisfied, the control logic can break down leading to mode destabilization, (3) two-loop control architectures that utilize small forward vanes with the aft tail provide highly attractive and feasible solutions to the longitudinal axis control challenges, and (4) closed-loop simulation sizing analyses indicate the baseline vane model utilized in this report is most likely oversized for normal loading conditions.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209528 , NAS 1.26:209528
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    Initial Piloted Simulation Evaluation of the Reference-H High-Speed Civil Transport Design During Takeoff and Recovery From Limit Flight Conditions (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: An initial assessment of a proposed High-Speed Civil Transport (HSCT) was conducted in the fall of 1995 at the NASA Langley Research Center. This configuration, known as the Industry Reference-H (Ref.-H), was designed by the Boeing Aircraft Company as part of their work in the High Speed Research program. It included a conventional tail, a cranked-arrow wing, four mixed-flow turbofan engines, and capacity for transporting approximately 300 passengers. The purpose of this assessment was to evaluate and quantify operational aspects of the Reference-H configuration from a pilot's perspective with the additional goal of identifying design strengths as well as any potential configuration deficiencies. This study was aimed at evaluating the Ref.-H configuration at many points of the aircraft's envelope to determine the suitability of the vehicle to accomplish typical mission profiles as well as emergency or envelope-limit conditions. Pilot-provided Cooper-Harper ratings and comments constituted the primary vehicle evaluation metric. The analysis included simulated real-time piloted evaluations, performed in a 6 degree of freedom motion base NASA Langley Visual-Motion Simulator, combined with extensive bath analysis. The assessment was performed using the third major release of the simulation data base (known as Ref.-H cycle 2B).
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209523 , NAS 1.26:209523
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    STS-103 Crew Interviews: Claude Nicollier (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-10
    Description: Live footage of a preflight interview with Mission Specialist Claude Nicollier is seen. The interview addresses many different questions including why Nicollier became an astronaut, the events that led to his interest, any role models that he had, and his inspiration. Other interesting information that this one-on-one interview discusses is an explanation of the why this required mission to service the Hubble Space Telescope must take place at such an early date, replacement of the gyroscopes, transistors, and computers. Also discussed are the Chandra X-Ray Astrophysics Facility, and a brief touch on Nicollier's responsibility during any of the given four space walks scheduled for this mission.
    Keywords: Aircraft Stability and Control
    Type: NONP-NASA-VT-1999213443 , JSC-1802G
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    Cellular Decomposition Based Hybrid-Hierarchical Control Systems with Applications to Flight Management Systems (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The work in this research project has been focused on the construction of a hierarchical hybrid control theory which is applicable to flight management systems. The motivation and underlying philosophical position for this work has been that the scale, inherent complexity and the large number of agents (aircraft) involved in an air traffic system imply that a hierarchical modelling and control methodology is required for its management and real time control. In the current work the complex discrete or continuous state space of a system with a small number of agents is aggregated in such a way that discrete (finite state machine or supervisory automaton) controlled dynamics are abstracted from the system's behaviour. High level control may then be either directly applied at this abstracted level, or, if this is in itself of significant complexity, further layers of abstractions may be created to produce a system with an acceptable degree of complexity at each level. By the nature of this construction, high level commands are necessarily realizable at lower levels in the system.
    Keywords: Aircraft Stability and Control
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    Updated Results of Deterministic Reconfigurable Control Design for the X-33 Vehicle (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-10
    Description: In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. An Off-line Nonlinear General Constrained Optimization approach was used for the reconfigurable X-33 control design method. Three examples of failure are shown using a high fidelity 6 DOF simulation (case 1: ascent with a left body flap jammed at 25 deg.; case 2: entry with a right inboard elevon jam at 25 deg. and case 3: landing (TAEM) (Terminal Area Energy Management) with a left rudder jam at -30 deg.) Failure comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 99-2934
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    The Development of the CONDUIT Advanced Control System Design and Evaluation Interface with a Case Study Application to an Advanced Fly by Wire Helicopter Design (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This report details the development and use of CONDUIT (Control Designer's Unified Interface). CONDUIT is a design tool created at Ames Research Center for the purpose of evaluating and optimizing aircraft control systems against handling qualities. Three detailed design problems addressing the RASCAL UH-60A Black Hawk are included in this report to show the application of CONDUIT to helicopter control system design.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1999-208763 , NAS 1.15:208763 , AFDD/TR-99-A-005 , A-99V-001
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    Flight-Determined, Subsonic, Lateral-Directional Stability and Control Derivatives of the Thrust-Vectoring F-18 High Angle of Attack Research Vehicle (HARV), and Comparisons to the Basic F-18 and Predicted Derivatives (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The subsonic, lateral-directional, stability and control derivatives of the thrust-vectoring F-1 8 High Angle of Attack Research Vehicle (HARV) are extracted from flight data using a maximum likelihood parameter identification technique. State noise is accounted for in the identification formulation and is used to model the uncommanded forcing functions caused by unsteady aerodynamics. Preprogrammed maneuvers provided independent control surface inputs, eliminating problems of identifiability related to correlations between the aircraft controls and states. The HARV derivatives are plotted as functions of angles of attack between 10deg and 70deg and compared to flight estimates from the basic F-18 aircraft and to predictions from ground and wind tunnel tests. Unlike maneuvers of the basic F-18 aircraft, the HARV maneuvers were very precise and repeatable, resulting in tightly clustered estimates with small uncertainty levels. Significant differences were found between flight and prediction; however, some of these differences may be attributed to differences in the range of sideslip or input amplitude over which a given derivative was evaluated, and to differences between the HARV external configuration and that of the basic F-18 aircraft, upon which most of the prediction was based. Some HARV derivative fairings have been adjusted using basic F-18 derivatives (with low uncertainties) to help account for differences in variable ranges and the lack of HARV maneuvers at certain angles of attack.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1999-206573 , NAS 1.60:206573 , H-2252
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    Analysis of Wind Tunnel Oscillatory Data of the X-31A Aircraft (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Wind tunnel oscillatory tests in pitch, roll, and yaw were performed on a 19%-scale model of the X-31A aircraft. These tests were used to study the aerodynamic characteristics of the X-31A in response to harmonic oscillations at six frequencies. In-phase and out-of-phase components of the aerodynamic coefficients were obtained over a range of angles of attack from 0 to 90 deg. To account for the effect of frequency on the data, mathematical models with unsteady terms were formulated by use of two different indicial functions. Data from a reduced set of frequencies were used to estimate model parameters, including steady-state static and dynamic stability derivatives. Both models showed good prediction capability and the ability to accurately fit the measured data. Estimated static stability derivatives compared well with those obtained from static wind tunnel tests. The roll and yaw rate derivative estimates were compared with rotary-balanced wind tunnel data and theoretical predictions. The estimates and theoretical predictions were in agreement at small angles of attack. The rotary-balance data showed, in general, acceptable agreement with the steady-state derivative estimates.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-208725 , NAS 1.26:208725
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    Technology Integration (Task 20) Aeroservoelastic Modeling and Design Studies (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The HSCT Flight Controls Group has developed longitudinal control laws, utilizing PTC aeroelastic flexible models to minimize aeroservoelastic interaction effects, for a number of flight conditions. The control law design process resulted in a higher order controller and utilized a large number of sensors distributed along the body for minimizing the flexibility effects. Processes were developed to implement these higher order control laws for performing the dynamic gust loads and flutter analyses. The processes and its validation were documented in Reference 2, for selected flight condition. The analytical results for additional flight conditions are presented in this document for further validation.
    Keywords: Aircraft Stability and Control
    Type: NF1676L-11122
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    Modeling the High Speed Research Cycle 2B Longitudinal Aerodynamic Database Using Multivariate Orthogonal Functions (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The data for longitudinal non-dimensional, aerodynamic coefficients in the High Speed Research Cycle 2B aerodynamic database were modeled using polynomial expressions identified with an orthogonal function modeling technique. The discrepancy between the tabular aerodynamic data and the polynomial models was tested and shown to be less than 15 percent for drag, lift, and pitching moment coefficients over the entire flight envelope. Most of this discrepancy was traced to smoothing local measurement noise and to the omission of mass case 5 data in the modeling process. A simulation check case showed that the polynomial models provided a compact and accurate representation of the nonlinear aerodynamic dependencies contained in the HSR Cycle 2B tabular aerodynamic database.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209525 , NAS 1.26:209525
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    Investigation of Inner Loop Flight Control Strategies for High-Speed Research (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: This report describes the activities and findings conducted under contract NAS1-19858 with NASA Langley Research Center. Subject matter is the investigation of suitable flight control design methodologies and solutions for large, flexible high-speed vehicles. Specifically, methodologies are to address the inner control loops used for stabilization and augmentation of a highly coupled airframe system possibly involving rigid-body motion, structural vibrations, unsteady aerodynamics, and actuator dynamics. Techniques considered in this body of work are primarily conventional-based, and the vehicle of interest is the High-Speed Civil Transport (HSCT). Major findings include 1) current aeroelastic vehicle modeling procedures require further emphasis and refinement, 2) traditional and nontraditional inner loop flight control strategies employing a single feedback loop do not appear sufficient for highly flexible HSCT class vehicles, 3) inner loop flight control systems will, in all likelihood, require multiple interacting feedback loops, and 4) Ref. H HSCT configuration presents major challenges to designing acceptable closed-loop flight dynamics.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209522 , NAS 1.26:209522
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    Modeling of Longitudinal Unsteady Aerodynamics of a Wing-Tail Combination (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Aerodynamic equations for the longitudinal motion of an aircraft with a horizontal tail were developed. In this development emphasis was given on obtaining model structure suitable for model identification from experimental data. The resulting aerodynamic models included unsteady effects in the form of linear indicial functions. These functions represented responses in the lift on the wing and tail alone, and interference between those two lifting surfaces. The effect of the wing on the tail was formulated for two different expressions concerning the downwash angle at the tail. The first expression used the Cowley-Glauert approximation known-as "lag-in-downwash," the second took into account growth of the wing circulation and delay in the development of the lift on the tail. Both approaches were demonstrated in two examples using the geometry of a fighter aircraft and a large transport. It was shown that the differences in the two downwash formulations would increase for an aircraft with long tail arm performing low-speed, rapid maneuvers.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1999-209547 , NAS 1.26:209547
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    Rotationally Adaptive Flight Test Surface (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Research on a new design of flutter exciter vane using adaptive materials was conducted. This novel design is based on all-moving aerodynamic surface technology and consists of a structurally stiff main spar, a series of piezoelectric actuator elements and an aerodynamic shell which is pivoted around the main spar. The work was built upon the current missile-type all-moving surface designs and change them so they are better suited for flutter excitation through the transonic flight regime. The first portion of research will be centered on aerodynamic and structural modeling of the system. USAF DatCom and vortex lattice codes was used to capture the fundamental aerodynamics of the vane. Finite element codes and laminated plate theory and virtual work analyses will be used to structurally model the aerodynamic vane and wing tip. Following the basic modeling, a flutter test vane was designed. Each component within the structure was designed to meet the design loads. After the design loads are met, then the deflections will be maximized and the internal structure will be laid out. In addition to the structure, a basic electrical control network will be designed which will be capable of driving a scaled exciter vane. The third and final stage of main investigation involved the fabrication of a 1/4 scale vane. This scaled vane was used to verify kinematics and structural mechanics theories on all-moving actuation. Following assembly, a series of bench tests was conducted to determine frequency response, electrical characteristics, mechanical and kinematic properties. Test results indicate peak-to-peak deflections of 1.1 deg with a corner frequency of just over 130 Hz.
    Keywords: Aircraft Stability and Control
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    Performance Modeling of a Pilot in a Free Flight Mode (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: Developments are being made that allow pilots to have more flexibility over the control of their aircraft. This new concept is called Free Flight. Free Flight strives to move the current air traffic system into an age where space technology is used to its fullest potential. Self-separation is one part of the Free Flight system. Self-separation provides pilots the opportunity to choose their own route to reach a specified destination provided that they maintain the 'minimum required separation distance between airplanes. In the event that pilots are unable to maintain separation, controllers will need to have the aircraft separation authority passed back to them. This situation is known as a procedural intervention point. This project attempted to examine and diagnose those particular situations in an effort to avoid reaching a procedural intervention point in the near future. Crews that reached procedural intervention points were compared with crews that made similar maneuver types in the same scenario, but did not reach procedural intervention points. Results showed that there were no significant differences between crews in a high-density acute angle flight conditions. However, significant differences in maneuver times, following the detection of an intruder aircraft and following the time the intruder aircraft came into view, were found in a low-density, acute angle scenario.
    Keywords: Aircraft Stability and Control
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    X-33 Attitude Control Using the XRS-2200 Linear Aerospike Engine (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: The Vehicle Control Systems Team at Marshall Space Flight Center, Structures and Dynamics Laboratory, Guidance and Control Systems Division is designing, under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control systems for the X-33 experimental vehicle. Test flights, while suborbital, will achieve sufficient altitudes and Mach numbers to test Single Stage To Orbit, Reusable Launch Vehicle technologies. Ascent flight control phase, the focus of this paper, begins at liftoff and ends at linear aerospike main engine cutoff (MECO). The X-33 attitude control system design is confronted by a myriad of design challenges: a short design cycle, the X-33 incremental test philosophy, the concurrent design philosophy chosen for the X-33 program, and the fact that the attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems. Additionally, however, and of special interest, the use of the linear aerospike engine is a departure from the gimbaled engines traditionally used for thrust vector control (TVC) in launch vehicles and poses certain design challenges. This paper discusses the unique problem of designing the X-33 attitude control system with the linear aerospike engine, requirements development, modeling and analyses that verify the design.
    Keywords: Aircraft Stability and Control
    Type: Joint Propulsion; Jun 20, 1999 - Jun 24, 1999; Los Angeles, CA; United States
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    Workstation-Based Simulation for Rapid Prototyping and Piloted Evaluation of Control System Designs (1998)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: The development and optimization of flight control systems for modem fixed- and rotary-. wing aircraft consume a significant portion of the overall time and cost of aircraft development. Substantial savings can be achieved if the time required to develop and flight test the control system, and the cost, is reduced. To bring about such reductions, software tools such as Matlab/Simulink are being used to readily implement block diagrams and rapidly evaluate the expected responses of the completed system. Moreover, tools such as CONDUIT (CONtrol Designer's Unified InTerface) have been developed that enable the controls engineers to optimize their control laws and ensure that all the relevant quantitative criteria are satisfied, all within a fully interactive, user friendly, unified software environment.
    Keywords: Aircraft Stability and Control
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    Implementation and Testing of Turbulence Models for the F18-HARV Simulation (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This report presents three methods of implementing the Dryden power spectral density model for atmospheric turbulence. Included are the equations which define the three methods and computer source code written in Advanced Continuous Simulation Language to implement the equations. Time-history plots and sample statistics of simulated turbulence results from executing the code in a test program are also presented. Power spectral densities were computed for sample sequences of turbulence and are plotted for comparison with the Dryden spectra. The three model implementations were installed in a nonlinear six-degree-of-freedom simulation of the High Alpha Research Vehicle airplane. Aircraft simulation responses to turbulence generated with the three implementations are presented as plots.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1998-206937 , NAS 1.26:206937
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    Optimization of Supersonic Transport Trajectories (1998)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: This paper develops a near-optimal guidance law for generating minimum fuel, time, or cost fixed-range trajectories for supersonic transport aircraft. The approach uses a choice of new state variables along with singular perturbation techniques to time-scale decouple the dynamic equations into multiple equations of single order (second order for the fast dynamics). Application of the maximum principle to each of the decoupled equations, as opposed to application to the original coupled equations, avoids the two point boundary value problem and transforms the problem from one of a functional optimization to one of multiple function optimizations. It is shown that such an approach produces well known aircraft performance results such as minimizing the Brequet factor for minimum fuel consumption and the energy climb path. Furthermore, the new state variables produce a consistent calculation of flight path angle along the trajectory, eliminating one of the deficiencies in the traditional energy state approximation. In addition, jumps in the energy climb path are smoothed out by integration of the original dynamic equations at constant load factor. Numerical results performed for a supersonic transport design show that a pushover dive followed by a pullout at nominal load factors are sufficient maneuvers to smooth the jump.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-112223 , NAS 1.15:112223 , A-98-09997
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    Robust Flutter Margin Analysis that Incorporates Flight Data (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An approach for computing worst-case flutter margins has been formulated in a robust stability framework. Uncertainty operators are included with a linear model to describe modeling errors and flight variations. The structured singular value, mu, computes a stability margin that directly accounts for these uncertainties. This approach introduces a new method of computing flutter margins and an associated new parameter for describing these margins. The mu margins are robust margins that indicate worst-case stability estimates with respect to the defined uncertainty. Worst-case flutter margins are computed for the F/A-18 Systems Research Aircraft using uncertainty sets generated by flight data analysis. The robust margins demonstrate flight conditions for flutter may lie closer to the flight envelope than previously estimated by p-k analysis.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-206543 , NAS 1.60:206543 , H-2209
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    Redesign of a Variable-Gain Output Feedback Longitudinal Controller Flown on the High-Alpha Research Vehicle (HARV) (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This paper describes a redesigned longitudinal controller that flew on the High-Alpha Research Vehicle (HARV) during calendar years (CY) 1995 and 1996. Linear models are developed for both the modified controller and a baseline controller that was flown in CY 1994. The modified controller was developed with three gain sets for flight evaluation, and several linear analysis results are shown comparing the gain sets. A Neal-Smith flying qualities analysis shows that performance for the low- and medium-gain sets is near the level 1 boundary, depending upon the bandwidth assumed, whereas the high-gain set indicates a sensitivity problem. A newly developed high-alpha Bode envelope criterion indicates that the control system gains may be slightly high, even for the low-gain set. A large motion-base simulator in the United Kingdom was used to evaluate the various controllers. Desired performance, which appeared to be satisfactory for flight, was generally met with both the low- and medium-gain sets. Both the high-gain set and the baseline controller were very sensitive, and it was easy to generate pilot-induced oscillation (PIO) in some of the target-tracking maneuvers. Flight target-tracking results varied from level 1 to level 3 and from no sensitivity to PIO. These results were related to pilot technique and whether actuator rate saturation was encountered.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-206938 , NAS 1.60:206938 , L-17640
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    A Worst-Case Approach for On-Line Flutter Prediction (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Worst-case flutter margins may be computed for a linear model with respect to a set of uncertainty operators using the structured singular value. This paper considers an on-line implementation to compute these robust margins in a flight test program. Uncertainty descriptions are updated at test points to account for unmodeled time-varying dynamics of the airplane by ensuring the robust model is not invalidated by measured flight data. Robust margins computed with respect to this uncertainty remain conservative to the changing dynamics throughout the flight. A simulation clearly demonstrates this method can improve the efficiency of flight testing by accurately predicting the flutter margin to improve safety while reducing the necessary flight time.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1997-207056 , NAS 1.15:207056
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    Distortion Tolerant Control Demonstrated in Flight (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Future aircraft turbine engines, both commercial and military, will have to be able to successfully accommodate expected increased levels of steady-state and dynamic engine-face distortion. Advanced tactical aircraft are likely to use thrust vectoring for enhanced aircraft maneuverability. As a result, the engines will see more extreme aircraft angle-of-attack alpha and sideslip beta levels than currently encountered with present-day aircraft. Also, the mixed-compression inlets needed for the High Speed Civil Transport (HSCT) will likely encounter disturbances similar to those seen by tactical aircraft, in addition to planar pulse, inlet buzz, and high distortion levels at low flight speed and off-design operation. The current approach of incorporating sufficient component design stall margin to tolerate these expected levels of distortion would result in significant performance penalties. The objective of NASA's High Stability Engine Control (HISTEC) program is to design, develop, and flight demonstrate an advanced, high-stability, integrated engine control system that uses measurement-based real-time estimates of distortion to enhance engine stability. The resulting distortion tolerant control adjusts the stall margin requirement online in real-time. This reduces the design stall margin requirement, with a corresponding increase in performance and decrease in fuel burn.
    Keywords: Aircraft Stability and Control
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    Computational Investigation and Validation of Twin-Tail Buffet Response Including Dynamics and Control (1998)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Multidisciplinary tools for prediction of single rectangular-tail buffet are extended to single swept-back-tail buffet in transonic-speed flow, and multidisciplinary tools for prediction and control of twin-tail buffet are developed and presented. The configuration model consists of a sharp-edged delta wing with single or twin tails that are oriented normal to the wing surface. The tails are treated as cantilevered beams fixed at the root and allowed to oscillate in both bending and torsion. This complex multidisciplinary problem is solved sequentially using three sets of equations on a dynamic single or multi-block grid structure. The first set is the unsteady, compressible, Reynolds-averaged Navier-Stokes equations which are used for obtaining the flow field vector and the aerodynamic loads on the tails. The Navier-Stokes equations are solved accurately in time using the implicit, upwind, flux-difference splitting, finite volume scheme. The second set is the coupled bending and torsion aeroelastic equations of cantilevered beams which are used for obtaining the bending and torsion deflections of the tails. The aeroelastic equations'are solved accurately in time using, a fifth-order-accurate Runge-Kutta scheme. The third set is the grid-displacement equations and the rigid-body dynamics equations, which are used for updating the grid coordinates due to the tail deflections and rigid-body motions. The tail-buffet phenomenon is predicted for highly-swept, single vertical tail placed at the plane of geometric symmetry, and for highly-swept, vertical twin tails placed at three different spanwise separation distances. The investigation demonstrates the effects of structural inertial coupling and uncoupling of the bending and torsion modes of vibration, spanwise positions of the twin-tail, angle of attack, and pitching and rolling dynamic motions of the configuration model on the tail buffet loading and response. The fundamental issue of twin-tail buffet alleviation is addressed using two active flow-control methods. These methods are the tangential leading-edge blowing and the flow suction from the leading-edge vortex cores along their paths. Qualitative and quantitative comparisons with the available experimental data are presented. The comparisons indicate that the present multidisciplinary aeroelastic analysis tools are robust, accurate and efficient.
    Keywords: Aircraft Stability and Control
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    Numerical Roll Reversal Predictor Corrector Aerocapture and Precision Landing Guidance Algorithms for the Mars Surveyor Program 2001 Missions (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes the development and evaluation of a numerical roll reversal predictor-corrector guidance algorithm for the atmospheric flight portion of the Mars Surveyor Program 2001 Orbiter and Lander missions. The Lander mission utilizes direct entry and has a demanding requirement to deploy its parachute within 10 km of the target deployment point. The Orbiter mission utilizes aerocapture to achieve a precise captured orbit with a single atmospheric pass. Detailed descriptions of these predictor-corrector algorithms are given. Also, results of three and six degree-of-freedom Monte Carlo simulations which include navigation, aerodynamics, mass properties and atmospheric density uncertainties are presented.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-4574 , AIAA Atmospheric Flight Mechanics Conference; Aug 10, 1998 - Aug 12, 1998; Boston, MA; United States
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    An Atmospheric Guidance Algorithm Testbed for the Mars Surveyor Program 2001 Orbiter and Lander (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: An Atmospheric Flight Team was formed by the Mars Surveyor Program '01 mission office to develop aerocapture and precision landing testbed simulations and candidate guidance algorithms. Three- and six-degree-of-freedom Mars atmospheric flight simulations have been developed for testing, evaluation, and analysis of candidate guidance algorithms for the Mars Surveyor Program 2001 Orbiter and Lander. These simulations are built around the Program to Optimize Simulated Trajectories. Subroutines were supplied by Atmospheric Flight Team members for modeling the Mars atmosphere, spacecraft control system, aeroshell aerodynamic characteristics, and other Mars 2001 mission specific models. This paper describes these models and their perturbations applied during Monte Carlo analyses to develop, test, and characterize candidate guidance algorithms.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-4569 , AIAA Atmospheric Flight Mechanics Conference; Aug 10, 1998 - Aug 12, 1998; Boston, MA; United States
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    Parameter Estimation of Actuators for Benchmark Active Control Technology (BACT) Wind Tunnel Model with Analysis of Wear and Aerodynamic Loading Effects (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This report describes the development of transfer function models for the trailing-edge and upper and lower spoiler actuators of the Benchmark Active Control Technology (BACT) wind tunnel model for application to control system analysis and design. A simple nonlinear least-squares parameter estimation approach is applied to determine transfer function parameters from frequency response data. Unconstrained quasi-Newton minimization of weighted frequency response error was employed to estimate the transfer function parameters. An analysis of the behavior of the actuators over time to assess the effects of wear and aerodynamic load by using the transfer function models is also presented. The frequency responses indicate consistent actuator behavior throughout the wind tunnel test and only slight degradation in effectiveness due to aerodynamic hinge loading. The resulting actuator models have been used in design, analysis, and simulation of controllers for the BACT to successfully suppress flutter over a wide range of conditions.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-208452 , NAS 1.15:208452 , L-17540 , AIAA Paper 96-3362 , Atmospheric Flight Mechanics Conference; Jul 29, 1996 - Jul 31, 1996; San Diego, CA; United States
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    Wavelet Filtering to Reduce Conservatism in Aeroservoelastic Robust Stability Margins (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wavelet analysis for filtering and system identification was used to improve the estimation of aeroservoelastic stability margins. The conservatism of the robust stability margins was reduced with parametric and nonparametric time-frequency analysis of flight data in the model validation process. Nonparametric wavelet processing of data was used to reduce the effects of external desirableness and unmodeled dynamics. Parametric estimates of modal stability were also extracted using the wavelet transform. Computation of robust stability margins for stability boundary prediction depends on uncertainty descriptions derived from the data for model validation. F-18 high Alpha Research Vehicle aeroservoelastic flight test data demonstrated improved robust stability prediction by extension of the stability boundary beyond the flight regime.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-206545 , NAS 1.15:206545 , H-2222 , AIAA Paper 98-1896 , Structures, Structural Dynamics and Materials Conference; Apr 20, 1998 - Apr 23, 1998; Long Beach, CA; United States
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    Flight Testing and Real-Time System Identification Analysis of a UH-60A Black Hawk Helicopter with an Instrumented External Sling Load (1998)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Helicopter external air transportation plays an important role in today's world. For both military and civilian helicopters, external sling load operations offer an efficient and expedient method of handling heavy, oversized cargo. With the ability to reach areas otherwise inaccessible by ground transportation, helicopter external load operations are conducted in industries such as logging, construction, and fire fighting, as well as in support of military tactical transport missions. Historically, helicopter and load combinations have been qualified through flight testing, requiring considerable time and cost. With advancements in simulation and flight test techniques there is potential to substantially reduce costs and increase the safety of helicopter sling load certification. Validated simulation tools make possible accurate prediction of operational flight characteristics before initial flight tests. Real time analysis of test data improves the safety and efficiency of the testing programs. To advance these concepts, the U.S. Army and NASA, in cooperation with the Israeli Air Force and Technion, under a Memorandum of Agreement, seek to develop and validate a numerical model of the UH-60 with sling load and demonstrate a method of near real time flight test analysis. This thesis presents results from flight tests of a U.S. Army Black Hawk helicopter with various external loads. Tests were conducted as the U.S. first phase of this MOA task. The primary load was a container express box (CONEX) which contained a compact instrumentation package. The flights covered the airspeed range from hover to 70 knots. Primary maneuvers were pitch and roll frequency sweeps, steps, and doublets. Results of the test determined the effect of the suspended load on both the aircraft's handling qualities and its control system's stability margins. Included were calculations of the stability characteristics of the load's pendular motion. Utilizing CIFER(R) software, a method for near-real time system identification was also demonstrated during the flight test program.
    Keywords: Aircraft Stability and Control
    Type: NASA/CR-1998-196710 , A-9809853 , NAS 1.26:196710
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    Modeling the Benchmark Active Control Technology Wind-Tunnel Model for Active Control Design Applications (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: This report describes the formulation of a model of the dynamic behavior of the Benchmark Active Controls Technology (BACT) wind tunnel model for active control design and analysis applications. The model is formed by combining the equations of motion for the BACT wind tunnel model with actuator models and a model of wind tunnel turbulence. The primary focus of this report is the development of the equations of motion from first principles by using Lagrange's equations and the principle of virtual work. A numerical form of the model is generated by making use of parameters obtained from both experiment and analysis. Comparisons between experimental and analytical data obtained from the numerical model show excellent agreement and suggest that simple coefficient-based aerodynamics are sufficient to accurately characterize the aeroelastic response of the BACT wind tunnel model. The equations of motion developed herein have been used to aid in the design and analysis of a number of flutter suppression controllers that have been successfully implemented.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-206270 , NAS 1.60:206270 , L-17625
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    Assessment Study of the State of the Art in Adaptive Control and its Applications to Aircraft Control (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Many papers relevant to reconfigurable flight control have appeared over the past fifteen years. In general these have consisted of theoretical issues, simulation experiments, and in some cases, actual flight tests. Results indicate that reconfiguration of flight controls is certainly feasible for a wide class of failures. However many of the proposed procedures although quite attractive, need further analytical and experimental studies for meaningful validation. Many procedures assume the availability of failure detection and identification logic that will supply adequately fast, the dynamics corresponding to the failed aircraft. This in general implies that the failure detection and fault identification logic must have access to all possible anticipated faults and the corresponding dynamical equations of motion. Unless some sort of explicit on line parameter identification is included, the computational demands could possibly be too excessive. This suggests the need for some form of adaptive control, either by itself as the prime procedure for control reconfiguration or in conjunction with the failure detection logic. If explicit or indirect adaptive control is used, then it is important that the identified models be such that the corresponding computed controls deliver adequate performance to the actual aircraft. Unknown changes in trim should be modelled, and parameter identification needs to be adequately insensitive to noise and at the same time capable of tracking abrupt changes. If however, both failure detection and system parameter identification turn out to be too time consuming in an emergency situation, then the concepts of direct adaptive control should be considered. If direct model reference adaptive control is to be used (on a linear model) with stability assurances, then a positive real or passivity condition needs to be satisfied for all possible configurations. This condition is often satisfied with a feedforward compensator around the plant. This compensator must be robustly designed such that the compensated plant satisfies the required positive real conditions over all expected parameter values. Furthermore, with the feedforward only around the plant, a nonzero (but bounded error) will exist in steady state between the plant and model outputs. This error can be removed by placing the compensator also in the reference model. Design of such a compensator should not be too difficult a problem since for flight control it is generally possible to feedback all the system states.
    Keywords: Aircraft Stability and Control
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  • 71
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    Effect of Configuration Pitching Motion on Twin Tail Buffet Response (1998)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The effect of dynamic pitch-up motion of delta wing on twin-tail buffet response is investigated. The computational model consists of a delta wing-twin tail configuration. The computations are carried out on a dynamic multi-block grid structure. This multidisciplinary problem is solved using three sets of equations which consists of the unsteady Navier-Stokes equations, the aeroelastic equations, and the grid displacement equations. The configuration is pitched-up from zero up to 60 deg. angle of attack, and the freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. With the twin tail fixed as rigid surfaces and with no-forced pitch-up motion, the problem is solved for the initial flow conditions. Next, the problem is solved for the twin-tail response for uncoupled bending and torsional vibrations due to the unsteady loads on the twin tail and due to the forced pitch-up motion. The dynamic pitch-up problem is also solved for the flow response with the twin tail kept rigid. The configuration is investigated for inboard position of the twin tail which corresponds to a separation distance between the twin tail of 33% wing chord. The computed results are compared with the available experimental data.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-0520 , Aerospace Sciences; Jan 12, 1998 - Jan 15, 1998; Reno, NV; United States
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    Effectiveness of Flow Control for Alleviation of Twin-Tail Buffet (1998)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Effectiveness of active flow control for twin- tail buffet alleviation is investigated. Tangen- tial leading-edge blowing (TLEB) and flow suction along the vortex cores (FSVC) of the lead- ing edges of the delta wing are used to delay the vortex breakdown flow upstream of the twin tail. The combined effect of the TLEB and FSVC is also investigated. A parametric study of the effects of the spanwise position of the suction tubes and volumetric suction flow rate on the twin-tail buffet response are also investigated. The TLEB moves the path of leading-edge vortices laterally towards the twin tail, which increases the aero- dynamic damping on the tails. The FSVC effectively delays the breakdown location at high angles of attack. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. This complex multidisciplinary problem is solved sequentially using three sets of equations for the fluid flow, aeroelastic response and grid deformation, on a dynamic multi-block grid structure. The computational model is pitched at 30 deg. angle of attack. The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. The model is investigated for the inboard position of the twin tails, which corresponds to a separation distance between the twin tails of 33% of the wing span.
    Keywords: Aircraft Stability and Control
    Type: Rept-985501 , 1998 World Aviation; Sep 28, 1998 - Sep 30, 1998; Anaheim, CA; United States
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  • 73
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    Effect of Apex Flap Deflection on Vertical Tail Buffeting (1998)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A computational study of the effect of vortex breakdown location on vertical tail buffeting is conducted. The position of the breakdown is modified by employing an apex flap deflected by an experimentally determined optimal angle. The delayed breakdown flow and buffeting response is then compared to the nominal undeflected case. This multidisciplinary problem is solved sequentially for the fluid flow, the elastic tail deformations and the grid displacements. The fluid flow is simulated by time accurately solving the unsteady, compressible, Reynolds-averaged Navier-Stokes equations using an implicit, upwind, flux-difference splitting finite volume scheme. The elastic vibrations of the tails are modeled by uncoupled bending and torsion beam equations. These equations are solved accurately in time using the Galerkin method and a five-stage Runge-Kutta-Verner scheme. The grid for the fluid dynamics calculations is continuously deformed using interpolation functions to disperse the displacements smoothly throughout the computational domain. An angle-of-attack of 35 deg.is chosen such that the wing primary-vortex cores experience vortex breakdown and the resulting turbulent wake flow impinges on tile vertical tails. The dimensions and material properties of the vertical tails are chosen such that the deflections are large enough to insure interaction with the flow, and the natural frequencies are high enough to facilitate a practical computational solution. Results are presented for a baseline uncontrolled buffeting case and a delayed breakdown case in which the apex flap has been deflected 15 deg. The flap was found to be very effective in delaying the breakdown, increasing the location from 50%c to 94%c, which resulted in a 6% increase in lift coefficient and pitching moment. However, the integrated buffet loads and tip responses were roughly equivalent for the two cases.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-0762 , Aerospace Sciences; Jan 12, 1998 - Jan 15, 1998; Reno, NV; United States
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  • 74
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    X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 98-4411 , GN and C Conference; Aug 11, 1998; Boston, MA; United States
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  • 75
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    Ascent, Transition, Entry, and Abort Guidance Algorithm Design for the X-33 Vehicle (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: One of the primary requirements for X-33 is that it be capable of flying autonomously. That is, onboard computers must be capable of commanding the entire flight from launch to landing, including cases where a single engine failure abort occurs. Guidance algorithms meeting these requirements have been tested in simulation and have been coded into prototype flight software. These algorithms must be sufficiently robust to account for vehicle and environmental dispersions, and must issue commands that result in the vehicle operating, within all constraints. Continual tests of these algorithms (and modifications as necessary) will occur over the next year as the X-33 nears its first flight. This paper describes the algorithms in use for X-33 ascent, transition, and entry flight, as well as for the powered phase of PowerPack-out (PPO) aborts (equivalent in thrust impact to losing an engine). All following discussion refers to these phases of flight when discussing guidance. The paper includes some trajectory results and results of dispersion analysis.
    Keywords: Aircraft Stability and Control
    Type: AIAA Paper 97-4409 , GN and C Conference; Aug 11, 1998; Boston, MA; United States
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  • 76
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    On-Line Robust Modal Stability Prediction using Wavelet Processing (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Wavelet analysis for filtering and system identification has been used to improve the estimation of aeroservoelastic stability margins. The conservatism of the robust stability margins is reduced with parametric and nonparametric time- frequency analysis of flight data in the model validation process. Nonparametric wavelet processing of data is used to reduce the effects of external disturbances and unmodeled dynamics. Parametric estimates of modal stability are also extracted using the wavelet transform. Computation of robust stability margins for stability boundary prediction depends on uncertainty descriptions derived from the data for model validation. The F-18 High Alpha Research Vehicle aeroservoelastic flight test data demonstrates improved robust stability prediction by extension of the stability boundary beyond the flight regime. Guidelines and computation times are presented to show the efficiency and practical aspects of these procedures for on-line implementation. Feasibility of the method is shown for processing flight data from time- varying nonstationary test points.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-206550 , H-2246 , NAS 1.15:206550 , ICAS-98-4,9,1 , Sep 14, 1998 - Sep 17, 1998; Melbourne; Australia
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    Flight Stability and Control and Performance Results from the Linear Aerospike SR-71 Experiment (LASRE) (1998)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The Linear Aerospike SR-71 Experiment (LASRE) is presently being conducted to test a 20-percent-scale version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single Stage to Orbit Technology Demonstrator Vehicle. The rocket engine was integrated into a lifting body configuration and mounted to the upper surface of an SR-71 aircraft. This paper presents stability and control results and performance results from the envelope expansion flight tests of the LASRE configuration up to Mach 1.8 and compares the results with wind tunnel predictions. Longitudinal stability and elevator control effectiveness were well-predicted from wind tunnel tests. Zero-lift pitching moment was mispredicted transonically. Directional stability, dihedral stability, and rudder effectiveness were overpredicted. The SR-71 handling qualities were never significantly impacted as a result of the missed predictions. Performance results confirmed the large amount of wind-tunnel-predicted transonic drag for the LASRE configuration. This drag increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could not be achieved on a hot day without depleting the available fuel.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-206565 , H-2276 , NAS 1.15:206565 , Atmosphere Flight Mechanics; Aug 10, 1998 - Aug 12, 1998; Boston, MA; United States
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    Accommodating Actuator Failures in Flight Control Systems (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: A technique for the design of flight control systems that can accommodate a set of actuator failures is presented. As employed herein, an actuator failure is defined as any change in the parametric model of the actuator which can adversely affect actuator performance. The technique is based upon the formulation of a fixed feedback topology which ensures at least stability in the presence of the failures in the set. The fixed compensation is obtained from a loop-shaping design procedure similar to Quantitative Feedback Theory and provides stability robustness in the presence of uncertainty in the vehicle dynamics caused by the failures. System adaptation to improve performance after actuator failure(s) occurs through a static gain adjustment in the compensator followed by modification of the system prefilter. Precise identification of the vehicle dynamics is unnecessary. Application to a single-input, single-output design using a simplified model of the longitudinal dynamics of the NASA High Angle of Attack Research Vehicle is discussed. Non-real time simulations of the system including a model of the pilot demonstrate the effectiveness and limitations of the approach.
    Keywords: Aircraft Stability and Control
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    A Control Law Design Method Facilitating Control Power, Robustness, Agility, and Flying Qualities Tradeoffs: CRAFT (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: A multi-input, multi-output control law design methodology, named "CRAFT", is presented. CRAFT stands for the design objectives addressed, namely, Control power, Robustness, Agility, and Flying Qualities Tradeoffs. The methodology makes use of control law design metrics from each of the four design objective areas. It combines eigenspace assignment, which allows for direct specification of eigenvalues and eigenvectors, with a graphical approach for representing the metrics that captures numerous design goals in one composite illustration. Sensitivity of the metrics to eigenspace choice is clearly displayed, enabling the designer to assess the cost of design tradeoffs. This approach enhances the designer's ability to make informed design tradeoffs and to reach effective final designs. An example of the CRAFT methodology applied to an advanced experimental fighter and discussion of associated design issues are provided.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-208463 , L-17571 , NAS 1.60:208463
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  • 80
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    Estimation of Aircraft Nonlinear Unsteady Parameters From Wind Tunnel Data (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: Aerodynamic equations were formulated for an aircraft in one-degree-of-freedom large amplitude motion about each of its body axes. The model formulation based on indicial functions separated the resulting aerodynamic forces and moments into static terms, purely rotary terms and unsteady terms. Model identification from experimental data combined stepwise regression and maximum likelihood estimation in a two-stage optimization algorithm that can identify the unsteady term and rotary term if necessary. The identification scheme was applied to oscillatory data in two examples. The model identified from experimental data fit the data well, however, some parameters were estimated with limited accuracy. The resulting model was a good predictor for oscillatory and ramp input data.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-208969 , NAS 1.15:208969 , L-17805
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    High-Alpha Research Vehicle Lateral-Directional Control Law Description, Analyses, and Simulation Results (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: This report contains a description of a lateral-directional control law designed for the NASA High-Alpha Research Vehicle (HARV). The HARV is a F/A-18 aircraft modified to include a research flight computer, spin chute, and thrust-vectoring in the pitch and yaw axes. Two separate design tools, CRAFT and Pseudo Controls, were integrated to synthesize the lateral-directional control law. This report contains a description of the lateral-directional control law, analyses, and nonlinear simulation (batch and piloted) results. Linear analysis results include closed-loop eigenvalues, stability margins, robustness to changes in various plant parameters, and servo-elastic frequency responses. Step time responses from nonlinear batch simulation are presented and compared to design guidelines. Piloted simulation task scenarios, task guidelines, and pilot subjective ratings for the various maneuvers are discussed. Linear analysis shows that the control law meets the stability margin guidelines and is robust to stability and control parameter changes. Nonlinear batch simulation analysis shows the control law exhibits good performance and meets most of the design guidelines over the entire range of angle-of-attack. This control law (designated NASA-1A) was flight tested during the Summer of 1994 at NASA Dryden Flight Research Center.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-208465 , NAS 1.60:208465 , L-17673
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    On the Numerical Formulation of Parametric Linear Fractional Transformation (LFT) Uncertainty Models for Multivariate Matrix Polynomial Problems (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: Robust control system analysis and design is based on an uncertainty description, called a linear fractional transformation (LFT), which separates the uncertain (or varying) part of the system from the nominal system. These models are also useful in the design of gain-scheduled control systems based on Linear Parameter Varying (LPV) methods. Low-order LFT models are difficult to form for problems involving nonlinear parameter variations. This paper presents a numerical computational method for constructing and LFT model for a given LPV model. The method is developed for multivariate polynomial problems, and uses simple matrix computations to obtain an exact low-order LFT representation of the given LPV system without the use of model reduction. Although the method is developed for multivariate polynomial problems, multivariate rational problems can also be solved using this method by reformulating the rational problem into a polynomial form.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-206939 , L-17720 , NAS 1.15:206939
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  • 83
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    A Method for Integrating Thrust-Vectoring and Actuated Forebody Strakes with Conventional Aerodynamic Controls on a High-Performance Fighter Airplane (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: A method, called pseudo controls, of integrating several airplane controls to achieve cooperative operation is presented. The method eliminates conflicting control motions, minimizes the number of feedback control gains, and reduces the complication of feedback gain schedules. The method is applied to the lateral/directional controls of a modified high-performance airplane. The airplane has a conventional set of aerodynamic controls, an experimental set of thrust-vectoring controls, and an experimental set of actuated forebody strakes. The experimental controls give the airplane additional control power for enhanced stability and maneuvering capabilities while flying over an expanded envelope, especially at high angles of attack. The flight controls are scheduled to generate independent body-axis control moments. These control moments are coordinated to produce stability-axis angular accelerations. Inertial coupling moments are compensated. Thrust-vectoring controls are engaged according to their effectiveness relative to that of the aerodynamic controls. Vane-relief logic removes steady and slowly varying commands from the thrust-vectoring controls to alleviate heating of the thrust turning devices. The actuated forebody strakes are engaged at high angles of attack. This report presents the forward-loop elements of a flight control system that positions the flight controls according to the desired stability-axis accelerations. This report does not include the generation of the required angular acceleration commands by means of pilot controls or the feedback of sensed airplane motions.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-1998-208464 , NAS 1.60:208464 , L-17627
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  • 84
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    Study of a Simulation Tool to Determine Achievable Control Dynamics and Control Power Requirements with Perfect Tracking (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: This paper contains a study of two methods for use in a generic nonlinear simulation tool that could be used to determine achievable control dynamics and control power requirements while performing perfect tracking maneuvers over the entire flight envelope. The two methods are NDI (nonlinear dynamic inversion) and the SOFFT(Stochastic Optimal Feedforward and Feedback Technology) feedforward control structure. Equivalent discrete and continuous SOFFT feedforward controllers have been developed. These equivalent forms clearly show that the closed-loop plant model loop is a plant inversion and is the same as the NDI formulation. The main difference is that the NDI formulation has a closed-loop controller structure whereas SOFFT uses an open-loop command model. Continuous, discrete, and hybrid controller structures have been developed and integrated into the formulation. Linear simulation results show that seven different configurations all give essentially the same response, with the NDI hybrid being slightly different. The SOFFT controller gave better tracking performance compared to the NDI controller when a nonlinear saturation element was added. Future plans include evaluation using a nonlinear simulation.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-208699 , L-17767 , NAS 1.15:208699
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    Flight-Time Identification of a UH-60A Helicopter and Slung Load (1998)
    In:  CASI
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    Publication Date: 2019-07-10
    Description: This paper describes a flight test demonstration of a system for identification of the stability and handling qualities parameters of a helicopter-slung load configuration simultaneously with flight testing, and the results obtained.Tests were conducted with a UH-60A Black Hawk at speeds from hover to 80 kts. The principal test load was an instrumented 8 x 6 x 6 ft cargo container. The identification used frequency domain analysis in the frequency range to 2 Hz, and focussed on the longitudinal and lateral control axes since these are the axes most affected by the load pendulum modes in the frequency range of interest for handling qualities. Results were computed for stability margins, handling qualities parameters and load pendulum stability. The computations took an average of 4 minutes before clearing the aircraft to the next test point. Important reductions in handling qualities were computed in some cases, depending, on control axis and load-slung combination. A database, including load dynamics measurements, was accumulated for subsequent simulation development and validation.
    Keywords: Aircraft Stability and Control
    Type: NASA/TM-1998-112231 , A-9810861 , NAS 1.15:112231 , USAATCOM-TR-98-A-004
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    Feedback Linearized Aircraft Control Using Dynamic Cell Structure (1998)
    In:  Other Sources
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    Publication Date: 2019-07-19
    Description: A Dynamic Cell Structure (DCS ) Neural Network was developed which learns a topology representing network (TRN) of F-15 aircraft aerodynamic stability and control derivatives. The network is combined with a feedback linearized tracking controller to produce a robust control architecture capable of handling multiple accident and off-nominal flight scenarios. This paper describes network and its performance for accident scenarios including differential stabilator lock, soft sensor failure, control, stability derivative variation, and turbulence.
    Keywords: Aircraft Stability and Control
    Type: WAC 98; May 10, 1998 - May 14, 1998; Anchorage, AK; United States
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  • 87
    Electronic Resource
    Electronic Resource
    Synthese, Eigenschaften und Struktur von LiAuI4 und KAuI4 mit einer Diskussion der kristallchemischen Verwandtschaft zwischen den Halogenoauraten RbAuCl4, AgAuCl4, RbAuBr4 und LiAuI4Professor Rudolf Hoppe zum 75. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1791-1795 
    Details
    ISSN: 0044-2313
    Keywords: Iodo Aurates(III) ; Synthesis ; Properties ; Crystal Structure ; Crystal Chemical Relationship ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Synthesis, Properties, and Structure of LiAuI4 and KAuI4 with a Discussion of the Crystal Chemical Relationship between the Halogenoaurates RbAuCl4, AgAuCl4, RbAuBr4 and LiAuI4The alkalimetal iodo aurates(III) MAuI4 (M = Li, K) are obtained in form of single crystals from MI, Au and I2 in a sealed glass ampoule by heating to 550°C and slow cooling to 300°C. KAuI4 crystallizes in the monoclinic space group P21/c with a = 968.6(4); b = 704.5(2), c = 1393.2(7) pm; β = 100.95(2)° and Z = 4. The crystal structure is built up from square planar AuI4- anions and K+ cations. The cations are coordinated by eight I atoms of neighbouring AuI4- anions with distances K—I between 350.0 and 369.6 pm. At 100°C KAuI4 is reduced to form K3Au3I8, which at 180°C decomposes to KI, Au and I2 LiAuI4 forms black, moisture sensitive needles, decomposing in the absence of iodine at 20°C to LiI, Au and I2. It crystallizes in a variant of the RbAuBr4 type structure with the space group P21/a and a = 1511.7(4); b = 433.9(4); c = 710.0(2) pm; β = 121.50(2)°; Z = 2. The crystal chemical relationship between the structures of RbAuCl4, RbAuBr4, AgAuCl4 and LiAuI4 is discussed.
    Notes: Die Alkalimetalliodoaurate(III) MAuI4 (M = Li, K) entstehen in Form von Einkristallen aus MI, Au und I2 in einer unter Vakuum abgeschmolzenen Glasampulle durch Erhitzen auf 550°C und langsames Abkühlen auf 300°C. KAuI4 kristallisiert monoklin in der Raumgruppe P21/c mit a = 968,6(4); b = 704,5(2), c = 1393,2(7) pm; β = 100,95(2)° und Z = 4. Die Kristallstruktur ist aus quadratisch-planaren AuI4--Anionen und K+-Kationen aufgebaut. Die Kationen sind von acht I-Atomen benachbarter AuI4--Ionen im Abstand von 350,0 bis 369,6 pm umgeben. KAuI4 zersetzt sich bei 100°C in K3Au3I8, das bei 180°C zu KI, Au und I2 zerfällt. LiAuI4 bildet schwarze, hydrolyseempfindliche Kristallnadeln, die sich außerhalb einer Iodatmosphäre bereits bei 20°C in LiI, Au und I2 zersetzen. Es kristallisiert in einer Strukturvariante des RbAuBr4-Typs mit der Raumgruppe P21/a und a = 1511,7(4); b = 433,9(4); c = 710,0(2) pm; β = 121,50(2)°; Z = 2. Die kristallchemische Verwandtschaft zwischen den Strukturen von RbAuCl4, RbAuBr4, AgAuCl4 und LiAuI4 wird diskutiert.
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    Lewis-Säure-Base-Reaktionen von Goldtrihalogeniden mit Bismuttrihalogeniden - Synthese und Kristallstrukturen von AuBiX6 (X = Cl, Br)Professor Rudolf Hoppe zum 75. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1810-1814 
    Details
    ISSN: 0044-2313
    Keywords: Gold trichloride ; gold tribromide ; bismuth trichloride ; bismuth tribromide ; solvothermal synthesis ; crystal structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Lewis-Acid-Base-Reactions of Gold Trihalides with Bismuth Trihalides - Synthesis and Structures of AuBiX6 (X = CI, Br)Gold trihalides AuX3 (X = Cl, Br) react with bismuth trihalides in sealed glass ampoules to the 1 : 1 adducts AuBiX6 (X = Cl, Br). AuBiCl6 is obtained by a chemical transport reaction at 220°C, whereas AuBiBr6 was synthesized by solvothermal reaction in SiBr4 at 150°C. Both compounds crystallize triclinic, space group P1, Z = 4. AuBiCl6; a = 698.3(4) pm; b = 1009.3(5) pm; c = 1381(1) pm; α = 104.98(5)°; β = 94.73(5)°; γ = 110.06(3)°; V = 867(1) · 106 pm3. AuBiBr6: a = 735.7(4) pm; b = 1055.7(5) pm; c = 1445(1) pm; α =104.88(5)°; β = 94.25(5)°; γ = 110.18(4)°; V =1001(1) ·106pm3. The structures are build formally of square-planar [AuX4]- and chains of edge-connected ([BiX4/2]+)n units. Since each Bi ion is surrounded by eight halogenide ions in a square-antiprismatic form, the structure can alternatively be described as consisting of chains of edge sharing ([BiX4X4/2]3-)n antiprisms connected by Au3+ ions.
    Notes: Goldtrihalogenide AuX3 (X = CI, Br) reagieren mit den entsprechenden Bismuttrihalogeniden BiX3 (X = Cl, Br) in geschlossenen Ampullen zu den 1 : 1-Addukten AuBiX6. AuBiCl6 entsteht bei 220°C unter den Bedingungen des chemischen Transportes, während AuBiBr6 wegen der geringen thermischen Stabilität von AuBr3 unter solvothermalen Bedingungen in SiBr4 als Lösungsmittel bei 150°C dargestellt wurde. Beide Verbindungen sind isotyp und kristallisieren triklin in der Raumgruppe P1, Z = 4, mit den Gitterkonstanten a = 698,3(4) pm; b = 1009,3(5) pm; c = 1381(1) pm; α = 104,98(5)°; β = 94,73(5)°; γ = 110,06(3)° für AuBiCl6 und a = 735,7(4) pm; b = 1055,7(5) pm; c = 1445(1) pm; α = 104,88(5)°; β = 94,25(5)°; γ = 110,18(4)° für AuBiBr6. Die Strukturen sind aus quadratisch-planaren [AuX4]--Ionen und Ketten aus kantenverknüpften ([BiX4/2]+)n-Einheiten aufgebaut. Da die Bi-Ionen von acht Halogenid-Ionen in Form eines quadratischen Antiprismas umgeben sind, kann die Struktur alternativ so beschrieben werden, daß sie aus Ketten kantenverknüpfter ([BiX4X4/2]3- )n-Antiprismen besteht, die über Au3+ -Ionen verbunden sind.
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    Ba11KX7O2 (X = P, As): Two Novel Zintl Phases with infinite chains of oxygen centered Ba6 octahedra, isolated X3- and dimeric X24- anionsProfessor Rudolf Hoppe zum 75. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1796-1802 
    Details
    ISSN: 0044-2313
    Keywords: Solid State Chemistry ; Crystal Structure ; Zintl Phases ; Oxides ; Phosphides ; Arsenides ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Ba11KX7O2 (X = P, As): Zwei neuartige Zintl-Phasen mit unendlichen Ketten sauerstoffzentrierter Ba6-Oktaeder, isolierten X3- und dimeren X24- AnionenReaktionen von „BaX“ (X = P, As) mit elementarem Ba, K und BaO in Tantalampullen bei 900-1000°C ergaben schwarze, sehr luft- und feuchtigkeitsempfindliche Kristalle von Ba11KP7O2 und der dazu isotypen Verbindung Ba11KAs7O2. welche durch EDX und Röntgenbeugung charakterisiert wurden (orthorhombisch, Fddd, Z = 8; a = 1069,9(1), b = 1514,3(2), c = 3164,6(4) pm und a = 1087,8(2), b = 1542,3(2), c = 3232,4(4) pm). In der Struktur liegen parallel zu [100] unendliche Zickzack-Ketten, ∞1[Ba4Ba2/2O], aus eckenverknüpften Ba6-Oktaedern vor, die durch O-Atome zentriert werden. Dazwischen befinden sich lineare Reihen von abwechselnd angeordneten isolierten X-Atomen und X2-Hanteln, welche die Oktaederketten zu Schichten parallel zu (001) miteinander verbinden. Während sich die X-Atome in einem verzerrten Ba8-Würfel befinden, zentrieren die X2-Hanteln einen Ba12-Polyeder, welcher aus einem Paar flächenverknüpfter, quadratischer Ba-Antiprismen aufgebaut ist. Hieraus ergibt sich die Abfolge Würfel-Antiprisma-Antiprisma-Würfel aus flächenverknüpften Ba-Polyedern. Zwischen den Schichten befinden sich weitere isolierte X-Atome und verknüpfen diese entlang [001] miteinander. Zwei Atompositionen sind statistisch mit Ba und K besetzt, so daß nach dem Zintl-Klemm-Konzept die Summenformel der Verbindungen auch durch Ba2+11K+X3-5(X2)4-O2-2 ausgedrückt werden kann.
    Notes: Reactions of “BaX” (X = P, As) with Ba, K and BaO in tantalum tubes at 900-1000°C yielded black, very air- and moisture-sensitive crystals of Ba11KP7O2 and isotypic Ba11KAs7O2 which were characterized by EDX and X-ray diffraction (orthorhombic, Fddd, Z = 8; a = 1069.9(1), b = 1514.3(2), c = 3164.6(4) pm and a = 1087.8(2), b = 1542.3(2), c = 3232.4(4) pm, respectively). The structure contains infinite zigzag chains, ∞1[Ba4Ba2/2O], of oxygen-centered, corner-sharing Ba6 octahedra along [100]. They are connected by linear strings built of alternating isolated X atoms and X2 dimers to form layers parallel to (001). While the isolated X atoms are surrounded by eight Ba forming a distorted cube, the X2 dimers center a Ba12 polyhedron which is comprised of a pair of face-sharing Ba square antiprisms. This results in a cube-antiprism-antiprism-cube sequence of face-sharing Ba polyhedra. Additional X atoms function as spacers between the layers and connect them along [001]. Two atom positions are statistically occupied by Ba and K, and the formula may be written as Ba2+11K+X3-5(X2)4-O2-2 according to the Zintl-Klemm concept.
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    Homoleptische Zinkamide: Der Übergang zu monomeren MolekülenProfessor Manfred Weidenbruch zum 60, Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1881-1884 
    Details
    ISSN: 0044-2313
    Keywords: Zinc amides ; monomeric zinc amides ; dimeric zinc amides ; X-ray structures ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Homoleptic Zinc Amides: Transition to Monomeric MoleculesZinc chloride reacts with the lithium salts of the bulky secondary amines HN(i-Bu)2 and HN(t-Bu)2 to form the corresponding homoleptic zinc compounds {Zn[N(i-Bu)2]2}2 (1) and Zn[N(t-Bu)2]2 (2). The NMR spectroscopic and mass spectrometric results as well as the molecular weight determinations and X-ray diffraction data are consistent with a dimeric structure of 1 and a monomeric structure of 2 in the gas phase, in solution and in the solid state.
    Notes: Zinkchlorid reagiert mit den Lithiumsalzen der sterisch anspruchsvollen sekundären Amine HN(i-Bu)2 und HN(t-Bu)2 zu den homoleptischen Zinkamiden {Zn[N(i-Bu)2]2}2 (1) und Zn[N(t-Bu)2]2 (2). Entsprechend der Ergebnisse von NMR-Spektroskopie, Massenspektrometrie, Molmassenbestimmung und Röntgenstrukturanalyse ist 1 dimer, 2 dagegen in der Gasphase, in Lösung und im Festkörper monomer.
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    Synthesis and reactions of {o-{[bis(dimethylamino)boryl]methyl}phenyl}diphenylphosphaneDedicated to Professor Welf Bronger on the Occasion of his 65th Birthday (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1093-1097 
    Details
    ISSN: 0044-2313
    Keywords: (Borylorganyl)phosphanes ; (Borylorganyl)phosphoniumsalt ; X-ray structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Synthese und Reaktionen von {o-{[Bis(dimethylamino)boryl]methyl}phenyl}diphenyl-phosphanDas neue α,ω-[Boryl(organyl)]phosphan o-Ph2PC6H4CH2 (NMe2)2 (10) wurde synthetisiert und spektroskopisch charakterisiert. Die Reaktivität dieser Verbindung, vor allem gegenüber Substitutionsreaktionen am Boratom war der Gegenstand unserer Untersuchungen. Verbindung 10 reagiert mit MeOH, BCl3 bzw. HCl zu den Verbindungen o-Ph2PC6H4CH2B(OMe)2 (10 a), o-Ph2PC6H4CH2BCl2 (10 c), o-Ph2PC6H4CH2BCl2 (HNMe2) (10 d) und o-Ph2P(HCl)C6H4CH2 BCl2(HNMe2) (10 e). Die Umsetzung von 10 a mit LiAlH4 führt zu o-Ph2PC6H4CH2BH2 (10 b). Verbindung 10 e wurde zusätzlich durch eine Röntgenstrukturanalyse charakterisiert.
    Notes: The new α,ω-[boryl(organyl)]phosphane o-Ph2PC6H4CH2B (NMe2)2 (10) was obtained in good yields from the reaction of CIB(NMe2)2 with o-Ph2PC6H4CH2Li(tmeda). Five derivatives of 10 were obtained by substituting the boron-bound amino groups by reactions with MeOH, BCl3, HCl, and LiAlH4, respectively, in particular, o-Ph2(HCl)PC6H4CH2BCl2 (HNMe2) (10 e) which shows a unique P—H—Cl—H—N unit. Compound 10 and its derivatives were characterized by multinuclear NMR methods, mass spectra, and elemental analyses. In addition, the structure of 10 e · 1.5 C6H6 was determined by single crystal X-ray diffraction.
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    Na2Mn(NH2)4: Ein neuer Schichtenstrukturtyp (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1108-1112 
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    ISSN: 0044-2313
    Keywords: Sodium tetra amido manganate(II) ; sodium manganese (II) amide ; X-ray structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Na2Mn(NH2)4: A New Type of Layered StructureThe structure of Na2Mn(NH2)4 was solved by X-ray single crystal data including H-positions: P21/c, Z = 4, a = 6.331(1) Å, b = 14.542(3) Å, c = 7.212(1) Å, β = 116.29(1)°, Z(F20 ≥ 3σ = (F20)) = 1343, Z(parameters) = 96, R/RW = 0.023/0.029.The compound crystallizes in a new type of structure. Within layered blocks the amide ions are arranged with the motif of a hexagonal closest packing of spheres. Within these blocks alternating layers contain sodium in all octahedral sites and manganese in an ordered way in a quarter of tetrahedral sites.
    Notes: Die Struktur von Na2Mn(NH2)4 wurde röntgenograhisch über Einkristalldaten einschließlich der H-Lagen bestimmt: P21/c, Z = 4, a = 6,331(1) Å, b = 14,542(3) Å, c = 7,212(1) Å, β = 116,29(1)°, Z(F0) mit (F0)2 ≥ 3σ(F0)2 = 1343, Z(Parameter) = 96, R/RW = 0,023/0,029.Die Verbindung kristallisiert in einem neuen Strukturtyp. Die Amidionen bilden innerhalb von Schichtenpaketen das Motiv einer hexagonal dichtesten Kugelpackung. In ihr sind in einzelnen Ebenen abwechselnd entweder alle Oktaederlücken durch Natrium oder geordnet 1/4 der Tetraederlücken durch Mangan besetzt.
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    Protonenlagen bei Kaliumtetraamidozinkat, K2Zn(NH2)4Professor Welf Bronger zum 65. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1103-1107 
    Details
    ISSN: 0044-2313
    Keywords: Potassium tetraamido zincate ; X-ray structure determination ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Positions of the Protons in Potassium Tetraamidozincate, K2Zn(NH2)4X-ray single crystal data for K2Zn(NH2)4 allowed the determination of the so far unknown positions of the protons: P1, Z = 2, a = 6.730(1) Å, b = 7.438(1) Å, c = 8.019(2) Å, α = 72.03(2)°, β = 84.45(2)°, γ = 63.82(1)°, Z(F0) with (F0)2 ≥ 3σ(F0)2 = 2166, Z(parameters) = 96, R/RW = 0.032/0.039.In the structure of K2Zn(NH2)2 the amide ions are nearly hexagonal close packed. One layer of octahedral holes parallel to (010) is fully occupied by potassium atoms and zinc is in an ordered way in a quarter of the tetrahedral holes of the next layer. The orientation of the protons of the amide ions is characteristic for this type of structure (filled up CdI2 type).
    Notes: Die bisher unbekannten Protonenlagen bei K2Zn(NH2)4 wurden röntgenographisch über Einkristalldaten bestimmt: P1, Z = 2, a = 6,730(1) Å, b = 7,438(1) Å, c = 8,019(2) Å, α = 72,03(2)°, β = 84,45(2)°, γ = 63,82(1)°, Z(F0) mit (F0)2 ≥ 3σ(F0)2 = 2166, Z(Parameter) = 96, R/RW = 0,032/0,039.Die Amidionen bilden das Motiv einer hexagonal dichtesten Kugelpackung. Darin sind in Ebenen parallel (010) abwechselnd alle Oktaederlücken durch Kalium oder geordnet 1/4 der Tetraederlücken durch Zink besetzt. Die Orientierung der Protonen an den Amidionen ist für diesen Strukturtyp (aufgefüllter CdI2-Typ) charakteristisch.
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    Derivate des Borabenzols. 25 [1]. Methylbis(1-methylboratabenzol)indium: Synthese und Struktur. Ein Beispiel für die Variabilität weicher molekularer Strukturparameter unter dem Einfluß der KristallgitterkräfteProfessor Welf Bronger zum 65. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1098-1102 
    Details
    ISSN: 0044-2313
    Keywords: Methylboratabenzene indium ; conformational variability ; automerization ; crystal packing effects ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Borabenzene Derivatives. 25 [1]. Methylbis(1-methylboratabenzene)indium: Synthesis and Structure. An Example for the Variability of Soft Structural Parameters in the Crystalline EnvironmentThe new boratabenzene derivative InMe(C5H5BMe)2 has been synthesized from 2-(Me3Sn)C5H5BMe and InMe3. The compound proves fluxional on the nmr time scale with average C2v symmetry in solution. An x-ray structure determination shows the existence of four molecules in the asymmetric unit differing mainly in the bonding geometry of the metal atoms with respect to the eight symmetrically independent boratabenzene ligands. Thus the dynamical behaviour in solution is reflected in the solid state by the soft structural parameters of the indium-ring coordination, with their variability exceeding the standard deviations of the corresponding individual observations by an order of magnitude.
    Notes: Ausgehend von 2-(Me3Sn)C5H5BMe und InMe3 wurde der neue Boratabenzol-Komplex InMe(C5H5BMe)2 erhalten. Die Verbindung zeigt in Lösung auf der NMR-Zeitskala eine fluktuierende Struktur mit effektiver C2v-Symmetrie. Die Kristallstrukturanalyse zeigt, daß im Festkörper vier Moleküle in der asymmetrischen Einheit vorliegen, die sich vorwiegend in der Bindung des Metallatoms an die insgesamt acht symmetrieunabhängigen Boratabenzol-Liganden unterscheiden. Das in Lösung dynamische Verhalten findet im Kristall seine Entsprechung in weichen Strukturparametern bezüglich der Indium-Ring-ligand-Koordination, deren Variabilität die Standardabweichungen der Einzelbeobachtung um eine Größenordnung übersteigt.
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    Derivate des Imidazols. XXIV [1].[Li12O2Cl2(ImN)8(THF)4] · 8 THF: Ein Peroxo-Komplex des Lithiums mit neuartiger KäfigstrukturProfessor Welter Siebert zum 60. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1121-1124 
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    ISSN: 0044-2313
    Keywords: Lithium ; imidazole derivatives ; peroxo complexes ; crystal structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Imidazole Derivatives. XXIV. [Li12O2Cl2(ImN)8(THF)4] · 8 THF: a Peroxo Lithium Fragment in a Novel Cage Structure1,3-dimethyl-2-iminoimidazoline (8, ImNH) reacts with methyl lithium to give [ImNLi]n (9). In tetrahydrofuran, crystals of C56H96Cl2Li12N24O6 · 8 C4H8O (10) are obtained. The structure of 10 consists of a Li12Cl2N8O2 core in which a peroxo unit is incorporated into a stack of ladder fragments. Over all, four tetrahydrofuran and eight imidazoline ligands are attached at the lithium and nitrogen atoms.
    Notes: 1,3-Dimethyl-2-iminoimidazolin (8, ImNH) reagiert mit Methyllithium in Diethylether zu [ImNLi]n (9). In Tetrahydrofuran werden Kristalle von [Li12O2Cl2(ImN)8(THF)4] · 8 THF (10) erhalten. In 10 liegt ein Li12Cl2N8O2-Käfig in Form gestapelter Leiterfragmente mit O22 als zentraler Einheit vor. Hieran sind über die Lithium- und Stickstoffatome insgesamt vier Tetrahydrofuran- und acht Imidazolin-Liganden koordiniert.
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    Ein ungewöhnlicher trimerer Bimetallkomplex mit dem Gerüst [Zr3(μ2-OH)3(μ3-O)Li5] durch Reaktion von Zirconiumorgano- und -hydrido-Komplexen mit Wasser (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1125-1130 
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    ISSN: 0044-2313
    Keywords: Zirconium ; Lithium ; Hydroxoligands ; Oxoligand ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: An Unusual Trimeric Bimetallic Li—Zr Complex with the Backbone [Zr3(μ2-OH)3(μ3-O)Li5] by Reaction of Zirconiumorgano and Hydrido Complexes with WaterThe reaction of compounds of the type [(LZr)(LiH)(L′)]n and [(LZr)(LiH)(L′)(alkyne)]n (L: 2,2′-Biphenolato-dianion, L′: thf, Bu3P, alkyne: Ph—C≡C—SiMe3, CH≡CH) with water at 0°C in a thf solution results in the formation of th trimeric bimetallic complex 8 [(L2Zr)3(μ2-OH)3(μ3-O)Li5(thf)8(H2O)5] in 50% yield. The X-ray analysis of 8 shows that a planar six-membered ring Zr3(μ2-OH)3 is formed. In the middle of this ring is a dianionic oxygen atom placed, coordinating to the three L2Zr centres in a planar μ3-coordination (bond angles 120,05μ). Five lithium ions stabilize the anionic backbone by bridging the biphenolato chelate ligands, which form seven-membered chelate rings with the atoms.1H-, 13C-, and 7Li-NMR spectra exhibit that the solid state structure remains unchanged in solution (thf).
    Notes: Die Reaktion von Verbindungen des Typs [(LZr)(LiH)(L′)]n und [(LZr)(LiH)(L′)(alkin)]n (L: 2,2′-Biphenolato-dianion, L′: thf, Bu3P, alkin: Ph—C≡C—SiMe3, CH≡CH) mit Wasser führt bei 0°C in thf in 50% Ausbeute zum trimeren Bimetallkomplex 8 [(L2Zr)3(μ2-OH)3(μ3-O) Li5(thf)8(H2O)5]. Die Kristallstrukturanalyse von 8 zeigt, daß ein planarer Zr3(μ2-OH)3-Sechsring gebildet wird. In seiner Mitte ist ein dianionisches Sauerstoffatom fixiert, das an den drei L2Zr-Zentren mit planarer μ3-Koordination gebunden ist (Bindungswinkel Zr—O—Zr 120,05°). Fünf Lithiumionen stabilisieren das anionische Gerüst durch Verbrückung der Biphenolatchelatliganden, die mit den Zr-Atomen Chelatsiebenringe ausbilden. 1H-, 13C- und 7Li-NMR-Spektren zeigen, daß in Lösung (thf) die Festkörperstruktur unverändert erhalten bleibt.
    Additional Material: 3 Ill.
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    URL: http://dx.doi.org/10.1002/zaac.19976230720
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  • 97
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    Synthesen und Strukturen der phosphor- und stickstoffverbrückten Übergangsmetallkomplexe [Pd(NPhPPh2)(PPh3)]2, [Pd(NPhPPh2)2 · Li(thf)]2, [Pd(NPhPPh2)Cl · Li(thf)3]2, [M(NPhPPh2)(HNPhPPh2)]2 (M = Pd, Pt), [M{Ph2P(NPh)2}2] (M = Co, Ni), [Ni(PPh2){Ph2P(NPh)2}]2 und [Ni2(PPh2)(NPhPPh2)(HNPhPPh2)3]Professor Wolfgang Beck zum 65. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1113-1120 
    Details
    ISSN: 0044-2313
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Syntheses and Structures of the Phosphorus and Nitrogenbridged Transition Metal Complexes [Pd(NPhPPh2)(PPh3)]2, [Pd(NPhPPh2)2 · Li(thf)]2, [Pd(NPhPPh2)Cl · Li(thf)3]2, [M(NPhPPh2)(HNPhPPh2)]2 (M=Pd, Pt), [M{Ph2P(NPh)2}2] (M=Co, Ni), [Ni(PPh2){Ph2P(NPh)2}]2 and [Ni2(PPh2)(NPhPPh2)(HNPhPPh2)3].From the reaction of LiNPhPPh2 with Palladium-Nickel- and Cobaltcomplexes, depending on the reaction conditions, different monomeric and dimeric complexes can be isolated. In these compounds the (NPhPPh2)--group acts as both a bridging and as a terminal ligand. [Pd(NPhPPh2)(PPh3)]2 (1), [Pd(NPhPPh2)2 · Li(thf)]2 (2) and [Pd(NPhPPh2)Cl · Li(thf)3]2 (3) are formed from the reaction of [PdCl2(PPh3)2] or [PdCl2(COD)] with LiNPhPPh2. In contrast to this from the reaction of Pd(Ac)2 and HNPhPPh2 (in the presence of zinc-dust) or [PtCl2(py)2] and LiNPhPPh2.
    Notes: Bei der Reaktion von LiNPhPPh2 mit Komplexen des Palladium, Cobalt und Nickel können je nach den Reaktionsbedingungen unterschiedliche einkernige und zweikernige Komplexe isoliert werden, in denen die (NPhPPh2)--Gruppe als Brücke bzw. endständiger Ligand vorliegt. [Pd(NPhPPh2)(PPh3)]2 (1), [Pd(NPhPPh2)2 · Li(thf)]2 (2) und [Pd(NPhPPh2)Cl · Li(thf)3]2 (3) werden bei der Umsetzung von [PdCl2(PPh3)2] bzw. [PdCl2(COD)] mit LiNPhPPh2 gebildet. Dagegen erhält man aus Pd(Ac)2 und HNPhPPh2 (in Gegenwart von Zinkstaub) bzw. [PtCl2(py)2] und LiNPhPPh2 die zweikernigen Komplexe [M(NPhPPh2)(HNPhPPh2)]2 (M = Pd: 4; M = Pt: 5). Die Umsetzung von [MCl2(PR3)2] (M = Ni: R = Ph, iPr; M = Co: R = Ph) führt dagegen zur Bildung der einkernigen Verbindungen [M{Ph2P(NPh)2}2] (M = Ni: 6; M = Co: 7) und des zweikernigen Komplexes [Ni(PPh2){Ph2P(NPh)2}]2 (8). Wird [NiCl2(PPh3)2] mit HNPhPPh2 in Gegenwart von Zinkstaub zur Reaktion gebracht, entsteht [Ni2(PPh2)(NPhPPh2)(HNPhPPh2)3] (9). Die Strukturen von 1-9 konnten mit Hilfe der Röntgenstrukturanalyse aufgeklärt werden. 1-9 besitzen folgende Gitterkonstanten und Raumgruppen:(1: Raumgruppe P1 (Nr. 2), Z = 2, a = 1193,0(4) pm, b = 1325,5(5) pm, c = 1447,2(7) pm, α = 97,01(3)°, β = 112,69(3)°, γ = 115,75(3)°; 2: Raumgruppe P1 (Nr. 2), Z = 2, a = 1264,1(13) pm, b = 1281,7(12) pm, c = 1448,0(2) pm, α = 113,01(6)°, β = 92,99(8)°, γ = 117,28(8)°; 3: Raumgruppe P1 (Nr.2) Z = 2, a = 1094,3(8) pm, b = 1197,5(12) pm, c = 1313,7(17) pm, α = 110,22(6)°, β = 101,33(6)°, γ = 106,45(5)°; 4: Raumgruppe P21/n (Nr. 14), Z = 4, a = 1490,1(4) pm, b = 1336,0(3) pm, c = 1746,9(6) pm, β = 97,52(3)°; 5: Raumgruppe P21/n (Nr. 14), Z = 4, a = 1486,7(2) pm, b = 1332,0(10) pm, c = 1749,0(2), β = 97,65(10)°, 6: Raumgruppe C2/c (Nr. 15), Z = 4, a = 1114,2(2) pm, b = 1722,7(3) pm, c = 2104,6(4) pm, β = 98,70(3); 7: Raumgruppe C2/c (Nr. 15), Z = 4, a = 1110,1(2) pm, b = 1714,3(3) pm, c = 2090,8(4) pm, β = 98,38(3)°; 8: Raumgruppe P1 (Nr. 2), Z = 2, a = 984,40(7) pm, b = 1189,7(8) pm, c = 1412,2(8) pm, γ = 101,49(5)°, β = 97,81(5), γ = 109,24(5); 9: Raumgruppe P1 (Nr. 2), Z = 2, a = 1402,1(9) pm, b = 1416,9(13) pm, c = 2230,0(2) pm, γ = 84,67(7)°, β = 84,98(6)°, γ = 65,72(6)°).
    Additional Material: 7 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/zaac.19976230718
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  • 98
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    The Structures and Characterization of Two New Oxides of the Sr4PtO6 Type: NaCa3IrO6 and NaCa3RuO6 (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1131-1134 
    Details
    ISSN: 0044-2313
    Keywords: Iridium ; ruthenium ; sodium ; calcium ; oxygen ; crystal structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Strukturen und Charakterisierung von zwei neuen Oxiden vom Sr4PtO6-Typ: NaCa3IrO6 und NaCa3RuO6NaCa3IrO6 (I) und NaCa3RuO6 (II) kristallisieren mit trigonaler (rhomboedrischer) Symmetrie in der Raumgruppe R3c, Z = 6, I: a = 9,272(3) Å, c = 11,214(1) Å, II: a = 9,244(3) Å, c = 11,201(1) Å. I und II sind isotyp zu Verbindungen vom Sr4PtO6-Typ. Die Strukturen wurden mittels Röntgen-Einkristallstrukturanalyse gelöst mit einem R-Wert von 0,032 und Rw = 0,039 für I bzw. R = 0,024 und Rw = 0,031 für II. Die Struktur besteht aus unendlichen Ketten von flächenverknüpften MO6-Oktaedern mit M=Ir bzw. Ru und trigonalen Prismen von NaO6. Die Ketten sind getrennt durch Calcium-Kationen.
    Notes: NaCa3IrO6 (I) and NaCa3RuO6 (II) crystallize with trigonal (rhombohedral) symmetry in the space group R3c, Z = 6, for I a = 9.272(3) Å, c = 11.214(1) Å; for II a = 9.244(3) Å, c = 11.201(1) Å. NaCa3IrO6 (I) and NaCa3RuO6 (II) are isotypic to compounds of the Sr4PtO6 structure type. The structures have been solved by means of single crystal X-ray diffraction data analysis with the reliability factors for I of R = 0.032 and Rw = 0.039; and the the reliability factors for II of R = 0.024 and Rw = 0.031. The structure consists of infinite chains of alternating face-sharing MO6 octahedra, where M=Ir or Ru, and NaO6 trigonal prisms. The chains are separated by the calcium cations.
    Additional Material: 3 Ill.
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    URL: http://dx.doi.org/10.1002/zaac.19976230721
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  • 99
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    Synthesis, characterization, and structures of 1-(9-Fluorenyl)-germatrane and 1-(Phenylacetylenyl)germatrane (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1144-1150 
    Details
    ISSN: 0044-2313
    Keywords: Trihalogenogermane ; Germatrane ; Transmetalation ; Crystal structure ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Darstellung, Charakterisierung und Strukturen von 1-(9-Fluorenyl)germatran and 1-(Phenylacetylenyl)germatranÜber die Synthesen der Titelverbindungen, nämlich N(CH2CH2O)3GeY (2 Y=Fluorenyl; 4 Y=PhC≡C) durch Reaktion von X3GeY (1 Y=Fluorenyl, X=Br; 5 Y=PhC≡C, X=Cl) mit N(CH2CH2OSnR3)3 (3 R=Et; 6 R=Bu) wird berichtet und die Darstellung der neuen Verbindung 1 wird beschrieben. Die Strukturen der Verbindungen wurden durch Elementaranalysen sowie mittels 1H and 13C NMR-Spektroskopie bestimmt, 2 und 4 massenspektrometrisch charakterisiert. 1, 2 und 4 wurden mit Röntgenbeugungsmethoden analysiert.
    Notes: Syntheses of the title compounds, viz. N(CH2CH2O)3GeY (2 Y=Fluorenyl; 4 Y=PhC≡C) by the reaction of X3GeY (1 Y=Fluorenyl, X=Br; 5 Y=PhC≡C, X=Cl) with N(CH2CH2OSnR3)3 (3 R=Et; 6 R=Bu) are reported including the preparation of the new compound 1. Identity and structures were established by elemental analyses, 1H and 13C NMR spectroscopy. 2 and 4 were characterized by mass spectrometry. Single crystal structures of 1, 2 and 4 were determined by X-ray diffraction methods.
    Additional Material: 3 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/zaac.19976230723
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  • 100
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    Synthesen und Kristallstrukturanalysen von Tetraalkylphosphonium-, -arsonium- und -stiboniumtriiodidenProfessor Walter Siebert zum 60. Geburtstag gewidmet (1997)
    Weinheim : Wiley-Blackwell
    Zeitschrift für anorganische Chemie 623 (1997), S. 1151-1156 
    Details
    ISSN: 0044-2313
    Keywords: Tetraalkyl phosphonium triiodide ; tetraalkyl arsonium triiodide ; tetraalkyl stibonium triiodide ; crystal structures ; Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Description / Table of Contents: Syntheses and Crystal Structure Analyses of Tetraalkyl Phosphonium, Arsonium, and Stibonium TriiodidesThe reaction of Me4EI (E=P, As), Me3EtSbI, Me2Et2SbI, MeEt3SbI, or Et4SbI with I2 in absence of solvent gives Me4PI3 (E=P, As), Me3EtSbI3, Me2Et2SbI3, MeEt3SbI3, or Et4SbI3. Me4SbI3 is formed in a reversible reaction by addition of I2 to (Me4Sb)3I8 or by reaction of a solution of Me4SbI in ethanol with I2 in benzene. The crystal structures of Me4EI3 (E=P, Sb), and Me3EtSbI3 and the syntheses of the novel compounds are reported.
    Notes: Die Reaktion von Me4EI (E=P, As), Me3EtSbI, Me2Et2SbI, MeEt3SbI oder Et4SbI mit I2 im Substanzgemisch ohne Lösungsmittel führt zu den entsprechenden Triiodiden Me4EI3 (E=P, As), Me3EtSbI3, Me2Et2SbI3, MeEt3SbI3 oder Et4SbI3. Me4SbI3 entsteht in reversibler Reaktion durch Zugabe von I2 zu (Me4Sb)3I8 oder durch Umsetzung einer ethanolischen Lösung von Me4SbI mit einer Lösung von Iod in Benzol. Über die Kristallstrukturen von Me4EI3 (E=P, Sb) und Me3EtSbI3 und die Synthese der neuen Verbindungen wird berichtet.
    Additional Material: 4 Ill.
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    URL: http://dx.doi.org/10.1002/zaac.19976230724
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