ALBERT

Description: A flight-test program to determine the flight characteristics of large jet transports equipped with powered-lift systems indicated the following results: Speed margins appear to be primarily related to power-on stall speeds. At these speed margins, the maneuver margins were adequate and did not appear to be a problem or limiting factor during powered-lift operation. No large detrimental effects on handling qualities were apparent but in some areas stability augmentation would be required to obtain satisfactory flight characteristics. Use of the powered-lift airplane would result in a sizable increase in the noise levels; these increases are primarily the result of higher engine power settings in the approach.

Description: During 5 1/2 years of flight research with a deflected-jet VTOL aircraft a 2 number of operational problems were experienced and investigated. These experiences can be grouped into two general categories: (1) the effects of the jet engine and its operation and (2) the restrictions imposed upon the pilot's operation due to reduced visual reference. This report is an examination of these problems and., in some cases, presents possible solutions to these problems.

Description: This paper covers the unique features and different modes of operation of V/STOL and STOL aircraft which result in new operating problems and serves to introduce the papers which deal in more detail with some of the more important V/STOL and STOL operating problems.

Description: Fixed-cockpit piloted simulator studies of delta-planform and variable-wing-sweep supersonic transport configurations are being conducted at the Ames Research Center to investigate the handling qualities and certification requirements related to the take-off maneuver. Validation of the simulation was achieved by duplicating the take-off certification program of a subsonic jet transport. Evaluation of the simulator was made by NASA pilots as well as company and FAA pilots involved in the actual certification flights of the airplane. The present paper is limited to a discussion of normal take-off, minimum control speed (ground), rotation characteristics, and initial climbout. Comparisons of the take-off characteristics are made between the supersonic transport and the current class of subsonic jet transports. Results indicate that minimum control speed (ground) characteristics are a function of thrust-weight ratio, the time provided for SST rotation should be at least as long as that for the subsonic jet transports, abused take-offs are more likely to result in tail scrapes, and climbout below the minimum drag speed requires that the pilot carefully monitor airspeed.

Description: Recent studies of NASA research related to aircraft operating problems on rough runways are presented. Some of these investigations were conducted cooperatively with the airport operators, with the Federal Aviation Agency, and with the U.S. Air Force. The studies show that criteria based on power spectral levels of runway-profile data are not sufficient to define acceptable levels of runway roughness from the piloting viewpoint. Because of the large variation in response characteristics between various types of aircraft, a runway may be acceptable for some aircraft and unacceptable for others. A criterion for roughness, therefore, should be expressed in terms of aircraft response - preferably, cockpit acceleration. A criterion suggested is that the maximum vertical acceleration in the cockpit should not exceed +/- 0.4 g for sections of the runway where precise aircraft control is required.

Description: Projections of technological advances to the year 2000 have suggested a number of possible conceptual types of air vehicles. For convenience, these have been grouped according to speed, increasing from below Mach 1 (including hovering flight) through various Mach ranges to orbital flight. The potential performance capabilities that could be realized for each of these types of aircraft provide exciting prospects for the years beyond 2000. It is clear that the greater design flexibilities provided by the projected technology advances will permit significant improvements in performance, economy, and safety and allow the design and development of aircraft systems that current technology will not support. In the following sections, each vehicle type is discussed, noting design features and significant advances. Where related aircraft exist, the significant economic and performance factors are compared. In addition, for each concept, the technology developments considered essential for the advanced concept are identified. The types of aircraft described are examples of what advances in the technology projected for the year 2000 could provide. They are by no means all that would be possible. The panel's views on the current state of knowledge of systems intergration as a discipline and the need for advancement in this area presented in the concluding section of this report.

Description: Over the last 50 years, aeronautical structures have evolved from the wood, fabric, and wire of low-speed biplanes to supersonic aircraft fabricated with advanced metallic and nonmetallic materials. The advances made in structures technology have had significant impact on aircraft design and performance. An example is the large weight reductions being realized from the utilization of composite materials for secondary and primary structures. However, structural advances have been evolutionary, not revolutionary. Through the year 2000, there are opportunities to obtain significant new advancements in structural technology. These advances could result in considerable performance and capability payoffs such as increased payload, range, speed, maneuverability, fuel efficiency, and safety through reductions in weight, increases in strength, and the ability to make structures pliable. Also, with new manufacturing processes it is possible that reductions in production costs will be realized. Some of the structural technology areas where future major advances could be made are adaptive structures, thermal structures, damage tolerant structures, propulsion system structures, and new structural concepts.

Description: A study has been made, using a piloted moving simulator, of the effects of the yaw-coupling parameters N(sub p) and N(sub delta(sub a) on the lateral-directional handling qualities of a large transport airplane at landing-approach airspeed. It is shown that the desirable combinations of these parameters tend to be more proverse when compared with values typical of current aircraft. Results of flight tests in a large variable-stability jet transport showed trends which were similar to those of the simulator data. Areas of minor disagreement, which were traced to differences in airplane geometry, indicate that pilot consciousness of side acceleration forces can be an important factor in handling qualities of future long-nosed transport aircraft.

Description: Although the major NASA research effort is directed toward XB-70-2, which will not enter its flight program until the summer of 1965, a limited amount of information is available from the early flights of the XB-70-1 airplane. Initial take-off and landing performance data have generally substantiated predictions and indicate no unforeseen problems for this class of vehicle. Vertical velocities at impact are of the same order of magnitude as those being experienced by present-day subsonic jets. The XB-70 distances from brake release to lift-off graphically illustrate the advantage of the increased thrust-weight ratio of the supersonic cruise vehicle. The landing loads are well within the design limits up to the highest vertical velocities encountered to date, and recorded data show the response at the pilot station to be somewhat greater than that recorded at the center of gravity. Persistent shaking has been encountered in flight at subsonic speed. The cause of the excitation is not known at present but the oscillation does not appear to be conventional buffeting. The oscillation occurrence drops off appreciably at supersonic speeds and can be correlated with atmospheric turbulence. The stability and control characteristics at subsonic speeds appear satisfactory with stability augmentation on and off. A longitudinal trim discrepancy from predictions has been noted in the transonic region which appears to be decreasing with increasing supersonic speed. The supersonic handling qualities are considered adequate with stability augmentation off; however, sensitive lateral control has resulted in small pilot-induced oscillations.

Description: An investigation of several factors which may contribute to the problem of piloting jet transport aircraft in heavy turbulence was conducted by using a piloted simulator that included the most significant airplane response and cockpit vibrations induced by rough air. Results indicated that the primary fuselage structural frequency contributed significantly to a distracting cockpit environment, and there was obtained evidence of severely reduced instrument flight proficiency during simulated maneuvering flight in heavy turbulence. It is concluded that the addition of similar rough-air response capabilities to training simulators would be of value in pilot indoctrination in turbulent-flight procedures.

Description: Recent results obtained from NASA V-G and VGH recorders installed on commercial turbine-powered transports have indicated that exceedances of placard speeds appear to have been significantly reduced since the placard speeds have been redefined and changes made in the aural warning. Oscillatory accelerations and unusual events, such as large or rapid departures from the planned flight profile, occur less frequently. Landing-impact accelerations are higher for turbine transports than for piston transports and vary with operator. The total in-flight acceleration experiences for turbine transports, however, are not significantly different from those for four-engine piston transports.

Description: The effects of surface temperature and imperfections on the drag of the supersonic transport are discussed. The relationships among surface temperature, emissivity, and skin friction are reviewed and the importance of manufacturing and maintenance imperfections is indicated.

Description: An investigation is currently under way to determine the operational practices and load experiences of general aviation aircraft performing five basic types of operations: twin-engine executive, single-engine executive, personal, instructional, and commercial survey. Limited data obtained to date from aircraft engaged in these operations indicate that aircraft are generally being operated within the limits to which they were designed.

Description: Studies of a number of STOL aircraft show that relatively high maximum lift-coefficients and large increases in lift due to power are within the present state of the art. With these lift characteristics, approach speeds of the order of 60 knots for aircraft of moderate wing loading can be realized. Full advantage of the STOL performance of aircraft such as those discussed herein may not be realized on a routine operational basis, however, without some form of damping augmentation system because of lateral-directional handling considerations, particularly for large aircraft operating under instrument flight conditions. Satisfactory characteristics can be obtained by use of only a servodriven rudder. Additional experience is needed to determine how the STOL aircraft is to be operated before more firm requirements for augmentation systems can be established.

Description: An analysis of the interaction of operational environment and aircraft characteristics of the supersonic transport (SST) in the areas of design-range and reserve-fuel requirements has been made. Design-range requirements are considered in relation to the effects of wind, temperature, flight-level assignment, and payload variation. An approach toward combining en route and holding reserve requirements while maintaining protection equivalent to that provided subsonic jet transport operations by the present civil air regulation en route plus holding reserves is given. This approach results in a savings in reserve fuel over that required by separate requirements.

Description: A flight program, utilizing a Convair 880 and a Boeing 720 airplane, was conducted in conjunction with wind-tunnel and simulator programs to study problems related to jet-transport upsets and operation in a turbulent environment. During the handling-qualities portion of the program the basic static stability of the airplanes was considered to be satisfactory and the lateral-directional damping was considered to be marginal without damper augmentation. An evaluation of the longitudinal control system indicated that this system can become marginal in effectiveness in the high Mach number and high dynamic-pressure range of the flight envelope. From the upset and recovery phase of the program it was apparent that retrimming the stabilizer and spoiler deployment were valuable tools in effecting a positive recovery; however, if these devices are to be used safely, it appears that a suitable g-meter should be provided in the cockpit because the high control forces in recovery tend to reduce the pilot's sensitivity to the actual acceleration loads. During the turbulence penetrations the pilot noted that the measured vibrations of 4 to 6 cps in the cockpit considerably disrupted their normal scan pattern and suggested that an improvement should be made in the seat cushion and restraint system. Also it was observed that the indicator needles on the flight instruments were quite stable in the turbulent environment.

Description: Several points of interest to the operator of aircraft are reviewed. Rates of crack propagation are shown to be high, and residual static strength is shown to decrease more significantly in high strength materials than in lower strength materials. Ground-air-ground and other negative loadings produce much more dn.mage than previously recognized or predicted. In first approximation, fatigue life is used in units corresponding to the number of flights, rather than to miles or hours as is commonly thought. Regular inspection for cracks is the only real way to assure safe operation of flight structures.

Description: Recent work on the traction of pneumatic tires on wet runways is discussed, and it is shown that a loss of tire traction adversely affects cross-wind landings. The effect of runway surface texture is discussed,, and a simple method for measuring surface texture is described. A preliminary correlation of tire traction with surface texture is shown. Results of work at Langley Research Center on the use of air jets to improve tire traction on wet or flooded runways indicate that this is a promising approach for alleviating the large losses in tire braking and sideways traction that occur when tire hydroplaning occurs on a flooded runway.

Description: Provides an overview of the X-48B prototype system flight test including vehicle characteristics and configuration. There are two X-48B Vehicles: the first, Vehicle 1, is the wind tunnel and flight test model. The second, Vehicle 2, provides the primary flight test. In mid-May 2006 the research team successfully completed 250 hours of wind tunnel tests on the X-48B Vehicle 1 at NASA's Langley Air Force Base. The prototype was then shipped to NASA's Dryden Flight Research Center at Edwards Air Force Base to serve as a backup to Vehicle 2, which is used for planned remotely piloted flight tests at Dryden.

Description: An overview of the technical development of hingeless and bearingless rotors is given, with emphasis on aeroelastic and aeromechanical stability characteristics. Important considerations for theoretical analysis are discussed. Theoretical and experimental investigations of isolated blade flap-lag and flap-lag-torsion stability, and coupled rotor-body aeromechanical stability are described. Physical interpretation and important rotor system design parameters are emphasized. An overview of bearingless rotor dynamics is also included.

Description: Raisbeck Mark Five Sabreliner 60A is a modernized, higher performance version of the Sabreliner business jet designed more than 20 years ago. Rockwell decided to update the design to incorporate NASA's computational fluid dynamics technology, thereby insuring continuing production of a competitive new airplane and protecting Sabreliner owners' investments by allowing them to convert to the update configuration. Redesign of the airplane involved extensive changes in aerodynamics structure and on-board systems. Key to Mark Five design is a new wing which has greater fuel capacity, increased span and improved airfoils to reduce fuel consumption.

Description: Rohr Industries, Inc. specializes in the manufacturing of nacelles, thrust reversers, and other engine components, has made use of the COSMIC program in nacelle work for aircraft like the McDonnell Douglas DC-10, the French/German Airbus and the Boeing 727, 737 and 747 jetliners. They also manufacture complete nacelles for military and business aircraft and are supplying nacelle components for the Boeing 757. The nacelle requires research and development for each type of airplane because of complex airflows around the engine inlet and high pressures on the nacelle skin. The use of the computer program defines the airflow field around turbofan engine nacelle inlet and cowls. Pressures on the nacelle skin are estimated for various flight conditions and structural integrity of the nacelle design. Rohr estimates that use of this program saved six man-months of programmer time necessary to develop alternative software.

Description: Farmers are increasingly turning to aerial applications of pesticides, fertilizers and other materials. Sometimes uneven distribution of the chemicals is caused by worn nozzles, improper alignment of spray nozzles or system leaks. If this happens, job must be redone with added expense to both the pilot and customer. Traditional pattern analysis techniques take days or weeks. Utilizing NASA's wind tunnel and computer validation technology, Dr. Roth, Oklahoma State University (OSU), developed a system for providing answers within minutes. Called the Rapid Distribution Pattern Evaluation System, the OSU system consists of a 100-foot measurement frame tied in to computerized analysis and readout equipment. System is mobile, delivered by trailer to airfields in agricultural areas where OSU conducts educational "fly-ins." A fly-in typically draws 50 to 100 aerial applicators, researchers, chemical suppliers and regulatory officials. An applicator can have his spray pattern checked. A computerized readout, available in five to 12 minutes, provides information for correcting shortcomings in the distribution pattern.

Description: We have developed a matrix calibration procedure that uniquely relates the electric fields measured at the aircraft with the external vector electric field and net aircraft charge. Our calibration method is being used with all of our aircraft/electric field sensing combinations and can be generalized to any reasonable combination of electric field measurements and aircraft. We determine a calibration matrix that represents the individual instrument responses to the external electric field. The aircraft geometry and configuration of field mills (FMs) uniquely define the matrix. The matrix can then be inverted to determine the external electric field and net aircraft charge from the FM outputs. A distinct advantage of the method is that if one or more FMs need to be eliminated or de-emphasized (for example, due to a malfunction), it is a simple matter to reinvert the matrix without the malfunctioning FMs. To demonstrate our calibration technique, we present data from several of our aircraft programs (ER-2, DC-8, Altus, Citation).

Description: This viewgraph presentation reviews the areas that Dryden Flight Research Center has set up for testing small Unmanned Aerial Systems (UAS). It also reviews the requirements and process to use an area for UAS test.

Description: This viewgraph presentation reviews Integrated Resilient Aircraft Control (IRAC) full scale flight tests.

Description: The primary objective of the UAVSAR Project is to develop a miniaturized polarimetric L-band synthetic aperture radar (SAR) for use on an unmanned aerial vehicle (UAV) or minimally piloted vehicle. Five Cycle 1 precision autopilot flights have been completed as of May 14, 2007. The first flight was open-loop controller, the second, third, fourth, and fifth flights were closed loop. The fifth flight demonstrated increasing duration within ten meter tube (approximately 90% of the time in the ten meter tube over a 200km course).

Description: In terms of technology, the X-43A/Hyper-X represented a singular milestone. After nearly a half century of high hopes, studies, wind tunnel tests, proposals, and canceled projects, a scramjet-powered vehicle had flown. The performance of the engine qualified the scramjet design tools and scaling laws. In turn, the theoretical calculations and ground testing could be used to design more advanced engine concepts. Just as important, both the scramjet and vehicle systems had successfully operated in the variable temperatures and densities of the atmosphere. The X-43A systems were able to maintain the exact flight conditions necessary for the scramjet to operate properly. Control deflections to correct the engine-induced moments were close to pre-flight predictions. When the unexpected occurred, such as when the vehicle pitched up during the cowl opening on the second flight, the control system was sufficiently designed to correct the situation. The airframe and wing structure, the thermal protection material, and the internal conditions of the X-43A performed largely as predicted. The HXLV thermal anomaly during the ascent on the third flight and "the Mach 8 unpleasantness" during the descent indicated that the HXLV and X-43A were not as resilient to aerodynamic heating as expected. The X-43A 's airframe drag and lift both were slightly higher than predicted, but still within preflight uncertainty predictions. The stability and control were as predicted, as was the boundary layer transition. The biggest aerodynamic worry before the flight was the separation of the HXLV and the X- 43A. After all was said and done, this went exactly as predicted, proving that non-symmetrical/high-dynamic pressure stage separations could be performed. This in turn meant that two-stage-to-orbit vehicles employing this technology were feasible. The Hyper-X program also served as a training ground for a new generation of scramjet and hypersonic researchers. This included both NASA and contractor personnel, providing them with experience in ground testing and component development; vehicle design, construction, integration, system checkout, and, ultimately, flight testing and data analysis. Additionally, researchers learned the practical details of running a project within finite budget and time limits, about the ambiguousness of risk assessment, and about the need to spend a significant amount of time and effort dealing with engineering problems, such as those with the FAS, that have nothing to do with the project's research goals. Finally, all those who worked on the X-43A project now know what it is like to spend years transforming an idea into a functional vehicle, only for it to be lost in a matter of seconds. And then to go through years of work to correct the problems, to face the possibility that still more might exist, and finally to savor the triumph of two successful flights. For those who will work on the hypersonic projects that emerge in coming years, these experiences may prove to be the most valuable of all.

Description: The NASA Dryden Flight Research Center, in partnership with the NASA Langley Research Center and industrial contractors, conducted the first flight tests of a supersonic combustion ramjet (scramjet) in 2004. This was a revolutionary airbreathing engine able to operate at speeds above Mach 5, which carries potential for both high-speed atmospheric flight and as a space launcher. For the Dryden engineers, the X-43 program was the culmination of a nearly 60-year history of flight research, going back to the early days of supersonic flight, and to rocket planes such as the X-1, D-558-II Skyrocket, and the X-15. For the propulsion community, it marked a turning point in a quest that had taken nearly as long. The scramjet engine did not arise from the work of a single individual or from a single technological breakthrough. It evolved instead from work under way on ramjets in the early 1950s, and from research programs at the National Advisory Committee for Aeronautics (NACA) Lewis Research Center, at the U.S. Army Aberdeen Proving Ground, and by the U.S. Navy. Studies developed in the course of these disparate projects raised the possibility of supersonic combustion. Many researchers had considered the notion impractical due to the difficulty of stabilizing a flame front in a supersonic airflow. NACA researchers at Lewis attempted to test the idea's feasibility by burning aluminum borohydride in a supersonic wind tunnel. Sustained burning was believed to have been observed at Mach 1.5, Mach 2, and Mach 3 for as long as two seconds.

Description: A viewgraph presentation of the Phoenix Missile Hypersonic Testbed (PMHT) is shown. The contents include: 1) Need and Goals; 2) Phoenix Missile Hypersonic Testbed; 3) PMHT Concept; 4) Development Objectives; 5) Possible Research Payloads; 6) Possible Research Program Participants; 7) PMHT Configuration; 8) AIM-54 Internal Hardware Schematic; 9) PMHT Configuration; 10) New Guidance and Armament Section Profiles; 11) Nomenclature; 12) PMHT Stack; 13) Systems Concept; 14) PMHT Preflight Activities; 15) Notional Ground Path; and 16) Sample Theoretical Trajectories.

Description: This viewgraph presentation describes the F-15 Intelligent Flight Control System (IFCS). The goals of this project include: 1) Demonstrate revolutionary control approaches that can efficiently optimize aircraft performance in both normal and failure conditions; and 2) Demonstrate advance neural network-based flight control technology for new aerospace systems designs.

Description: The Unmanned Aerial Vehicle Synthetic Aperture Radar (UAVSAR) project began as an Instrument Incubator Program (IIP) out of the NASA ESTO Program Office. After a year of study JPL presented to NASA an instrument concept that could be accommodated on the desired class of platforms, that would meet the original IIP science and instrument objectives and could be expanded to meet future airborne radar science needs. The UAVSAR project is a four year program consisting of a 3 year phase in which the radar system is designed and fabricated, the platform is modified, radar is installed on the aircraft and an initial flight testing program is begun. The last year of the program is designed to collect repeat pass data, to improve system robustness and to validate that the scientific objectives of the sensor are being met.

Description: A general overview of the Orion abort flight test is presented. The contents include: 1) Abort Flight Test Project Overview; 2) DFRC Exploration Mission Directorate; 3) Abort Flight Test; 4) Flight Test Configurations; 5) Flight Test Vehicle Engineering Office; 6) DFRC FTA Scope; 7) Flight Test Operations; 8) DFRC Ops Support; 9) Launch Facilities; and 10) Scope of Launch Abort Flight Test

Description: A viewgraph presentation of flight tests performed on the F/A active aeroelastic wing airplane is shown. The topics include: 1) F/A-18 AAW Airplane; 2) F/A-18 AAW Control Surfaces; 3) Flight Test Background; 4) Roll Control Effectiveness Regions; 5) AAW Design Test Points; 6) AAW Phase I Test Maneuvers; 7) OBES Pitch Doublets; 8) OBES Roll Doublets; 9) AAW Aileron Flexibility; 10) Phase I - Lessons Learned; 11) Control Law Development and Verification & Validation Testing; 12) AAW Phase II RFCS Envelopes; 13) AAW 1-g Phase II Flight Test; 14) Region I - Subsonic 1-g Rolls; 15) Region I - Subsonic 1-g 360 Roll; 16) Region II - Supersonic 1-g Rolls; 17) Region II - Supersonic 1-g 360 Roll; 18) Region III - Subsonic 1-g Rolls; 19) Roll Axis HOS/LOS Comparison Region II - Supersonic (open-loop); 20) Roll Axis HOS/LOS Comparison Region II - Supersonic (closed-loop); 21) AAW Phase II Elevated-g Flight Test; 22) Region I - Subsonic 4-g RPO; and 23) Phase II - Lessons Learned

Description: Flight testing of the 14ft span CloudSwift UAV was conducted during the summer of 2005. Test maneuvers included aircraft checkout, Piccolo gain tuning, FTS range tests, and thermal soaring research flights.

Description: A stiffened composite panel has been designed that is representative of the fuselage structure of existing wide bodied aircraft. The panel is a minimum weight design, based on the current level of technology and realistic loads and criteria. Several different stiffener configurations were investigated in the optimization process. The final configuration is an all graphite epoxy J-stiffened design in which the skin between adjacent stiffeners is permitted to buckle under design loads. Fail-safe concepts typically employed in metallic fuselage structure have been incorporated in the design. A conservative approach has been used with regard to structural details such as skin frame and stringer frame attachments and other areas where sufficient design data was not available.

Description: A preliminary methodology was obtained for the design optimization of a subsonic aircraft by coupling NASA Langley Research Center s Flight Optimization System (FLOPS) with NASA Glenn Research Center s design optimization testbed (COMETBOARDS with regression and neural network analysis approximators). The aircraft modeled can carry 200 passengers at a cruise speed of Mach 0.85 over a range of 2500 n mi and can operate on standard 6000-ft takeoff and landing runways. The design simulation was extended to evaluate the optimal airframe and engine parameters for the subsonic aircraft to operate on nonstandard runways. Regression and neural network approximators were used to examine aircraft operation on runways ranging in length from 4500 to 7500 ft.

Description: A computational investigation is underway at the NASA Glenn Research Center to determine the aerodynamic performance of subsonic scarf inlets. These inlets are characterized as being longer over the lower portion of the inlet, as shown in the preceding figure. One of the key variables being investigated in the research is the circumferential extent of the longer portion of the inlet. It shows two specific geometries that are being examined: one in which the length of the inlet transitions from long-to-short over the full 180 deg. from bottom to top, and a second in which the length transitions over 67.5 deg.

Description: With the advent of ultrahigh-bypass engines, the space available for passive acoustic treatment is becoming more limited, whereas noise regulations are becoming more stringent. Active noise control (ANC) holds promise as a solution to this problem. It uses secondary (added) noise sources to reduce or eliminate the offending noise radiation. The first active noise control test on the low-speed fan test bed was a General Electric Company system designed to control either the exhaust or inlet fan tone. This system consists of a "ring source," an induct array of error microphones, and a control computer. Fan tone noise propagates in a duct in the form of spinning waves. These waves are detected by the microphone array, and the computer identifies their spinning structure. The computer then controls the "ring source" to generate waves that have the same spinning structure and amplitude, but 180 out of phase with the fan noise. This computer generated tone cancels the fan tone before it radiates from the duct and is heard in the far field. The "ring source" used in these tests is a cylindrical array of 16 flat-plate acoustic radiators that are driven by thin piezoceramic sheets bonded to their back surfaces. The resulting source can produce spinning waves up to mode 7 at levels high enough to cancel the fan tone. The control software is flexible enough to work on spinning mode orders from -6 to 6. In this test, the fan was configured to produce a tone of order 6. The complete modal (spinning and radial) structure of the tones was measured with two builtin sets of rotating microphone rakes. These rakes provide a measurement of the system performance independent from the control system error microphones. In addition, the far-field noise was measured with a semicircular array of 28 microphones. This test represents the first in a series of tests that demonstrate different active noise control concepts, each on a progressively more complicated modal structure. The tests are in preparation for a demonstration on a flight-type engine.

Description: Pressure-distribution measurements have been made on the fus elage of the Bell X- 1 research airplane. Data are presented for angles of attack from 2 deg. to 8 deg. during pull-ups at Mach numbers of about 0.78, 0.85, 0.88, and 1.02. The results of the investigation indicated that a large portion of the load carried by the fuselage was in the vicinity of the wing and may be attributed to wing-to-fuselage carryover. The presence of the wing from the 41 to 60 percent fuselage stations influenced the fuselage pressures from about 30 to 65 percent fuselage length at Mach numbers of approximat ely 0.78, 0.85, and 0.88, and from about 35 to 80 percent fuselage length at a Mach number of approximately 1.02. The fuselage contributed about 20 percent of the total airplane normal-force coefficient. The center of pressure of the fuselage load throughout the tests was located from 41 to 51 percent fuselage length, which corresponds to the forward half of the wing root-chord location.

Description: A method has been proposed for predicting the effect of a rapid blade-pitch increase on the thrust and induced-velocity response of a helicopter rotor. General equations have been derived for the ensuing motion of the helicopter. These equations yield time histories of thrust, induced velocity, and helicopter vertical velocity for given rates of blade-pitch-angle changes and given rotor-angular-velocity time histories. The results of the method have been compared with experimental results obtained with a rotor mounted on the Langley helicopter test tower. The calculated and experimental results are in good agreement, although, in general, the calculated thrust-coefficient overshoots are about 10 percent greater than those obtained experimentally.

Description: A cascade of 65-(12)10 compressor blades was tested at one geometric setting over a range of inlet Mach number from 0.12 to 0.89. Two groups of data are presented and compared: the first from the cascade operating conventionally with no boundary-layer control, and the second with the boundary layer controlled by a combination of upstream slot suction and porous-wall suction at the blade tips. A criterion for two-dimensionality was used to specify the degree of boundary-layer control by suction to be applied. The data are presented and an analysis is made to show the effect of Mach number on turning angle, blade wake, pressure distribution about the blade profile and static-pressure rise. The influence of boundary-layer control on these parameters as well as on the secondary losses is illustrated. A system of correlating the measured static-pressure rise through the cascade with the theoretical isentropic values is presented which gives good agreement with the data. The pressure distribution about the blade profile for an inlet Mach number of 0.21 is corrected with the Prandtl-Glauert, Karman-Tsien, and vector-mean velocity - contraction coefficient compressibility correction factors to inlet Mach numbers of 0.6 and 0.7. The resulting curves are compared with the experimental pressure distributions for inlet Mach numbers of 0.6 and 0.7 so that the validity of applying the three corrections can be evaluated.

Description: The empirical relation between the induced velocity, thrust, and rate of vertical descent of a helicopter rotor was calculated from wind tunnel force tests on four model rotors by the application of blade-element theory to the measured values of the thrust, torque, blade angle, and equivalent free-stream rate of descent. The model tests covered the useful range of C(sub t)/sigma(sub e) (where C(sub t) is the thrust coefficient and sigma(sub e) is the effective solidity) and the range of vertical descent from hovering to descent velocities slightly greater than those for autorotation. The three bladed models, each of which had an effective solidity of 0.05 and NACA 0015 blade airfoil sections, were as follows: (1) constant-chord, untwisted blades of 3-ft radius; (2) untwisted blades of 3-ft radius having a 3/1 taper; (3) constant-chord blades of 3-ft radius having a linear twist of 12 degrees (washout) from axis of rotation to tip; and (4) constant-chord, untwisted blades of 2-ft radius. Because of the incorporation of a correction for blade dynamic twist and the use of a method of measuring the approximate equivalent free-stream velocity, it is believed that the data obtained from this program are more applicable to free-flight calculations than the data from previous model tests.

Description: The autorotative performance of an assumed helicopter was studied to determine the effect of inoperative jet units located at the rotor-blade tip on the helicopter rate of descent. For a representative ramjet design, the effect of the jet drag is to increase the minimum rate of descent of the helicopter from about 1,OO feet per minute to 3,700 feet per minute when the rotor is operating at a tip speed of approximately 600 feet per second. The effect is less if the rotor operates at lower tip speeds, but the rotor kinetic energy and the stall margin available for the landing maneuver are then reduced. Power-off rates of descent of pulse-jet helicopters would be expected to be less than those of ramjet. helicopters because pulse jets of current design appear to have greater ratios of net power-on thrust to power-off, drag than currently designed rain jets. Iii order to obtain greater accuracy in studies of autorotative performance, calculations in'volving high power-off rates of descent should include the weight-supporting effect of the fuselage parasite-drag force and the fact that the rotor thrust does not equal the weight of the helicopter.

Description: Jet noise and flow field were measured to follow up on observations made by Professor D. Papamoschou of the University of California at Irvine (NASA Grant NAG3-2345). When a dual-stream coannular nozzle was arranged non-concentrically, noise was attenuated significantly on the side where the annulus was thicker. A similar observation was also made in reference 2. The practical significance is obvious. If the bypass flow of a jet exhaust in flight could be diverted to form a thicker layer underneath, then less noise would be heard by an observer on the ground. In view of the current emphasis on jet noise abatement, researchers felt that the effect deserved further attention. This prompted an experiment to confirm the phenomenon in a larger facility and to obtain flow-field data to advance understanding of the mechanism.

Description: The Parametric Inlet is an innovative concept for the inlet of a gas-turbine propulsion system for supersonic aircraft. The concept approaches the performance of past inlet concepts, but with less mechanical complexity, lower weight, and greater aerodynamic stability and safety. Potential applications include supersonic cruise aircraft and missiles. The Parametric Inlet uses tailored surfaces to turn the incoming supersonic flow inward toward an axis of symmetry. The terminal shock spans the opening of the subsonic diffuser leading to the engine. The external cowl area is smaller, which reduces cowl drag. The use of only external supersonic compression avoids inlet unstart--an unsafe shock instability present in previous inlet designs that use internal supersonic compression. This eliminates the need for complex mechanical systems to control unstart, which reduces weight. The conceptual design was conceived by TechLand Research, Inc. (North Olmsted, OH), which received funding through NASA s Small-Business Innovation Research program. The Boeing Company (Seattle, WA) also participated in the conceptual design. The NASA Glenn Research Center became involved starting with the preliminary design of a model for testing in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10 10 SWT). The inlet was sized for a speed of Mach 2.35 while matching requirements of an existing cold pipe used in previous inlet tests. The parametric aspects of the model included interchangeable components for different cowl lip, throat slot, and sidewall leading-edge shapes and different vortex generator configurations. Glenn researchers used computational fluid dynamics (CFD) tools for three-dimensional, turbulent flow analysis to further refine the aerodynamic design.

Description: NASA Glenn Research Center s Flywheel Development Team designed, built, and successfully operated the new G2 flywheel to 41,000 rpm on September 2, 2004. This work was supported by the Aerospace Flywheel Technology Program--a NASA Office of Aerospace Technology ETC Program funded by the Energetics Project. The work was performed by a team of civil servants, contractors, and grantees managed by Glenn s Electrical Systems Development Branch, Structural Mechanics and Dynamics Branch, and Space Power & Propulsion Test Engineering Branch. The G2 flywheel was designed to be a low-cost modular testbed for flywheel system integration and component demonstrations.

Description: An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

Description: The powered-lift Channel Wing concept has been combined with pneumatic Circulation Control aerodynamic and propulsive technology to generate a Pneumatic Channel Wing (PCW) configuration intended to have Super-STOL or VSTOL capability while eliminating many of the operational problem areas of the original Channel Wing vehicle. Wind-tunnel development and evaluations of a PCW powered model conducted at Georgia Tech Research Institute (GTRI) have shown substantial lift capabilities for the blown configuration (CL values of 10 to 11). Variation in blowing of the channel was shown to be more efficient than variation in propeller thrust in terms of lift generation. Also revealed was the ability to operate unstalled at very high angles of attack of 40 deg - 45 deg, or to achieve very high lift at much lower angle of attack to increase visibility and controllability. In order to provide greater flexibility in Super-STOL takeoffs and landings, the blown model also displayed the ability to interchange thrust and drag by varying blowing without any moving parts. A preliminary design study of this pneumatic vehicle based on the two technologies integrated into a simple Pneumatic Channel Wing configuration showed very strong Super-STOL potential. This paper presents these experimental results, discusses variations in the configuration geometry under development, and addresses additional considerations to extend this integrated technology to advanced design studies of PCW-type vehicles.

Description: NASA s Vehicle Systems Program is investing in aeronautics technology development across six vehicle sectors, in order to improve future air travel. These vehicle sectors include subsonic commercial transports, supersonic vehicles, Uninhabited Aerial Vehicles (UAVs), Extreme Short Takeoff and Landing (ESTOL) vehicles, Rotorcraft, and Personal Air Vehicles (PAVs). While the subsonic transport is firmly established in U.S. markets, the other vehicle sectors have not developed a sufficient technology or regulatory state to permit widespread, practical use. The PAV sector has legacy products in the General Aviation (GA) market, but currently only accounts for negligible revenue miles, sales, or market share of personal travel. In order for PAV s to ever capture a significant market, these small aircraft require technologies that permit them to be less costly, environmentally acceptable, safer, easier to operate, more efficient, and less dependent on large support infrastructures.

Description: Circulation Control technologies have been around for 65 years, and have been successfully demonstrated in laboratories and flight vehicles alike, yet there are few production aircraft flying today that implement these advances. Circulation Control techniques may have been overlooked due to perceived unfavorable trade offs of mass flow, pitching moment, cruise drag, noise, etc. Improvements in certain aspects of Circulation Control technology are the focus of this paper. This report will describe airfoil and blown high lift concepts that also address cruise drag reduction and reductions in mass flow through the use of pulsed pneumatic blowing on a Coanda surface. Pulsed concepts demonstrate significant reductions in mass flow requirements cor Circulation Control, as well as cruise drag concepts that equal or exceed conventional airfoil systems.

Description: The vision of NASA s Dryden Flight Research Center is to "fly what others only imagine." Its mission is to advance technology and science through flight. Objectives supporting the mission include performing flight research and technology integration to revolutionize aviation and pioneer aerospace technology, validating space exploration concepts, conducting airborne remote sensing and science missions, and supporting operations of the Space Shuttle and the International Space Station. A significant focus of effort in recent years has been on Unmanned Aircraft Systems (UAS), both in support of the Airborne Science Program and as research vehicles to advance the state of the art in UAS. Additionally, the Center has used its piloted aircraft in support of UAS technology development. In order to facilitate greater access to the UAS expertise that exists at the Center, that expertise has been organized around three major capabilities. The first is access to high-altitude, long-endurance UAS. The second is the establishment of a test range for small UAS. The third is safety case assessment support.

Description: The Joint Unmanned Combat Air Systems (J-UCAS) program is a collaborative effort between the Defense Advanced Research Project Agency (DARPA), the US Air Force (USAF) and the US Navy (USN). Together they have reviewed X-45A flight test site processes and personnel as part of a system demonstration program for the UCAV-ATD Flight Test Program. The goal was to provide a disciplined controlled process for system integration and testing and demonstration flight tests. NASA's Dryden Flight Research Center (DFRC) acted as the project manager during this effort and was tasked with the responsibilities of range and ground safety, the provision of flight test support and infrastructure and the monitoring of technical and engineering tasks. DFRC also contributed their engineering knowledge through their contributions in the areas of autonomous ground taxi control development, structural dynamics testing and analysis and the provision of other flight test support including telemetry data, tracking radars, and communications and control support equipment. The Air Force Flight Test Center acted at the Deputy Project Manager in this effort and was responsible for the provision of system safety support and airfield management and air traffic control services, among other supporting roles. The T-33 served as a J-UCAS surrogate aircraft and demonstrated flight characteristics similar to that of the the X-45A. The surrogate served as a significant risk reduction resource providing mission planning verification, range safety mission assessment and team training, among other contributions.

Description: Objectives: a) Safely and efficiently perform structural load tests on NAVAIR E-2C aircraft to calibrate strain gage instrumentation installed by NAVAIR; b) Collect load test data and derive loads equations for use in NAVAIR flight tests; and c) Assist flight test team with use of loads equations measurements at PAX River.

Description: This presentation reports the results of the NASA/DARPA automatic probe and drogue refueling flight test. The program met several of its objectives including the design, development and successful testing of a prototype system to autonomously perform probes to drogue refueling; demonstrated acquisition and tracking capability of the video tracking system; demonstrated autonomous rendezvous capability; demonstrated ability to plug in a turn; and, demonstrated ability to plug in mild turbulence.

Description: A general overview of the vortex induced performance benefits of Dryden's DC-8 and F-18 trail aircraft is shown. The status of the mission and safety hazards are also presented.

Description: The use of face gears in an advanced rotorcraft transmission design was first proposed by the McDonnell Douglas Helicopter Company during their contracted effort with the U.S. Army under the Advanced Rotorcraft Transmission (ART) program. Face gears would be used to turn the corner between the horizontal gas turbine engine and the vertical output rotor shaft--a function currently done by spiral bevel gears. This novel gearing arrangement would substantially lower the drive system weight partly because a face gear mesh would be used to split the input power between two output gears. However, the use of face gears and their ability to operate successfully at the speeds and loads required for an aerospace environment was unknown. Therefore a proof-of-concept phase with an existing test stand at the NASA Lewis Research Center was pursued. Hardware was designed that could be tested in Lewis' Spiral Bevel Gear Test Rig. The initial testing indicated that the face gear mesh was a feasible design that could be used at high speeds and load. Surface pitting fatigue was the typical failure mode, and that could lead to tooth fracture. An interim project was conducted to see if slight modifications to the gear tooth geometry or an alternative heat treating process could overcome the surface fatigue problems. From the initial and interim tests, it was apparent that for the surface fatigue problems to be overcome the manufacturing process used for this component would have to be developed to the level used for spiral bevel gears. The current state of the art for face gear manufacturing required using less than optimal gear materials and manufacturing techniques because the surface of the tooth form does not receive final finishing after heat treatment as it does for spiral bevel gears. This resulted in less than desirable surface hardness and manufacturing tolerances. An Advanced Research and Projects Agency (ARPA) Technology Reinvestment Project has been funded to investigate the effects of manufacturing process improvements on the operating characteristics of face gears. The program is being conducted with McDonnell Douglas Helicopter Co., Lucas Western Inc., the University of Illinois at Chicago, and a NASA/U.S. Army team. The goal of the project is develop the grinding process, experimentally verify the improvement in face gear fatigue life, and conduct a full-scale helicopter transmission test. The theory and methodology to grind face gears has been completed, and manufacture of the test hardware is ongoing. Experimental verification on test hardware is scheduled to begin in fiscal 1996.

Description: This slide presentation reviews the work of the Experimental Capabilities Supersonic project, that is being reorganized into Flight Research and Validation. The work of Experimental Capabilities Project in FY '09 is reviewed, and the specific centers that is assigned to do the work is given. The portfolio of the newly formed Flight Research and Validation (FRV) group is also reviewed. The various projects for FY '10 for the FRV are detailed. These projects include: Eagle Probe, Channeled Centerbody Inlet Experiment (CCIE), Supersonic Boundary layer Transition test (SBLT), Aero-elastic Test Wing-2 (ATW-2), G-V External Vision Systems (G5 XVS), Air-to-Air Schlieren (A2A), In Flight Background Oriented Schlieren (BOS), Dynamic Inertia Measurement Technique (DIM), and Advanced In-Flight IR Thermography (AIR-T).

Description: This paper presents the progress in the development of a low-cost change-detection system. This system is being developed to provide users with the ability to use a low-cost unmanned aerial vehicle (UAV) and image processing system that can detect changes in specific fixed ground locations using video provided by an autonomous UAV. The results of field experiments conducted with the US Army at Ft. A.P.Hill are presented.

Description: The pressure recovery of incoming cooling air and the drag associated with engine cooling of a typical general aviation twin-engine aircraft was Investigated experimentally. The semispan model was mounted vertically in the 40 x 80-Foot Wind Tunnel at Ames Research Center. The propeller was driven by an electric motor to provide thrust with low vibration levels for the cold-now configuration. It was found that the propeller slip-stream reduces the frontal air spillage around the blunt nacelle shape. Consequently, this slip-stream effect promotes flow reattachment at the rear section of the engine nacelle and improves inlet pressure recovery. These effects are most pronounced at high angles of attack; that is, climb condition. For the cruise condition those improvements were more moderate.

Description: This paper presents an overview of years of sensor system development and application for aerospace systems. The emphasis of this work is on developing advanced capabilities for measurement and control of aeropropulsion and crew vehicle systems as well as monitoring the safety of those systems. Specific areas of work include chemical species sensors, thin film thermocouples and strain gages, heat flux gages, fuel gages, SiC based electronic devices and sensors, space qualified electronics, and MicroElectroMechanical Systems (MEMS) as well as integrated and multifunctional sensor systems. Each sensor type has its own technical challenges related to integration and reliability in a given application. The general approach has been to develop base sensor technology using microfabrication techniques, integrate sensors with "smart" hardware and software, and demonstrate those systems in a range of aerospace applications. Descriptions of the sensor elements, their integration into sensors systems, and examples of sensor system applications will be discussed. Finally, suggestions related to the future of sensor technology will be given. It is concluded that smart micro/nano sensor technology can revolutionize aerospace applications, but significant challenges exist in maturing the technology and demonstrating its value in real-life applications.

Description: The Structures & Materials Discipline within the NASA Subsonic Rotary Wing Project is focused on developing rotorcraft technologies. The technologies being developed are within the task areas of: 5.1.1 Life Prediction Methods for Engine Structures & Components 5.1.2 Erosion Resistant Coatings for Improved Turbine Blade Life 5.2.1 Crashworthiness 5.2.2 Methods for Prediction of Fatigue Damage & Self Healing 5.3.1 Propulsion High Temperature Materials 5.3.2 Lightweight Structures and Noise Integration The presentation will discuss rotorcraft specific technical challenges and needs as well as details of the work being conducted in the six task areas.

Description: In order to develop the capability to evaluate control system technologies, NASA Ames Research Center (Ames) began a test program to build a Hover Test Vehicle (HTV) - a ground-based simulated flight vehicle. The HTV would integrate simulated propulsion, avionics, and sensors into a simulated flight structure, and fly that test vehicle in terrestrial conditions intended to simulate a flight environment, in particular for attitude control. The ultimate purpose of the effort at Ames is to determine whether the low-cost hardware and flight software techniques are viable for future low cost missions. To enable these engineering goals, the project sought to develop a team, processes and procedures capable of developing, building and operating a fully functioning vehicle including propulsion, GN&C, structure, power and diagnostic sub-systems, through the development of the simulated vehicle.

Description: The Five-Axis, Three-Magnetic-Bearing Dynamic Spin Rig is an apparatus for vibration testing of turbomachine blades in a vacuum at rotational speeds from 0 to 40,000 rpm. This rig includes (1) a vertically oriented shaft on which is mounted an assembly comprising a rotor holding the blades to be tested, (2) two actively controlled heteropolar radial magnetic bearings at opposite ends of the shaft, and (3) an actively controlled magnetic thrust bearing at the upper end of the shaft. This rig is a more capable successor to a prior apparatus, denoted the Dynamic Spin Rig (DSR), that included a vertically oriented shaft with a mechanical thrust bearing at the upper end and a single actively controlled heteropolar radial magnetic bearing at the lower end.

Description: A short report discusses selected aspects of the development of the University of Florida micro-aerial vehicle (UFMAV) basically, a miniature airplane that has a flexible wing and is representative of a new class of airplanes that would operate autonomously or under remote control and be used for surveillance and/or scientific observation. The flexibility of the wing is to be optimized such that passive deformation of the wing in the presence of aerodynamic disturbances would reduce the overall response of the airplane to disturbances, thereby rendering the airplane more stable as an observation platform. The aspect of the development emphasized in the report is that of computational simulation of dynamics of the UFMAV in flight, for the purpose of generating mathematical models for use in designing control systems for the airplane. The simulations are performed by use of data from a wind-tunnel test of the airplane in combination with commercial software, in which are codified a standard set of equations of motion of an airplane, and a set of mathematical routines to compute trim conditions and extract linear state space models.

Description: The damping in roll and rolling effectiveness of two models of a missile having cruciform, triangular, interdigitated wings and tails have been determined through a Mach number range of 0.8 to 1.8 by utilizing rocket-propelled test vehicles. Results indicate that the damping in roll was relatively constant over the Mach umber range investigated. The rolling effectiveness was essentially constant at low supersonic speeds and increased with increasing mach numbers in excess of 1.4 over the Mach number range investigated. Aeroelastic effects increase the rolling-effectiveness parameters pb/2V divided by delta and decrease both the rolling-moment coefficient due to wing deflection and the damping-in-roll coefficient.

Description: The Rake Airflow Gage Experiment was flown on the Propulsion Flight Test Fixture at NASA Dryden Flight Research Center using one of Dryden s F-15B research testbed aircraft. Propulsion Flight Test Fixture is a modular, pylon-based platform for flight testing propulsion system components, such as the Channeled Centerbody Inlet Experiment, an innovative, variable-geometry, mixed compression supersonic inlet under development at NASA Dryden. The objective of this flight test was to ascertain the flowfield angularity and local Mach number profile of the aerodynamic interface plane that is defined by the planned location of the tip of the inlet centerbody. Knowledge of the flowfield characteristics at this location underneath will be essential to computational modeling of the new inlet as well as future propulsion systems flight testing using the test fixture. This paper describes the preparation for and execution of the flight test, as well as results and validation of the algorithm used to calculate local Mach number and angularity from the rake's pressure measurements.

Description: Airframe structural morphing technologies designed to mitigate sonic boom strength are being developed by Gulfstream Aerospace Corporation (GAC). Among these technologies is a concept in which an aircraft's frontend would be extended prior to supersonic acceleration. This morphing would effectively lengthen the vehicle, reducing peak sonic boom amplitude, but is also expected to partition the otherwise strong bow shock into a series of reduced-strength, non-coalescing shocklets. This combination of boom shaping techniques is predicted to transform the classic, high-impulse N-wave pattern typically generated by an aircraft traveling at supersonic speed into a signature more closely resembling a sinusoidal wave with a greatly reduced perceived loudness. 'QuietSpike' is GAC's nomenclature for its recently patented front-end vehicle morphing arrangement. The ability of Quiet Spike to effectively shape a vehicle's far- field sonic boom signature is highly dependent on the area distribution characteristics of the aircraft. The full aeroacoustic benefits of front-end morphing at farfield are only possible when the QuietSpike article and vehicle configuration are designed in consideration of each other. Adding QuietSpike technology to the airframe of an existing, non-boom-optimized supersonic vehicle is unlikely to result in an improved far-field signature due to the generally over-powering influence of wing- and inlet-generated shocks. Therefore, it is generally recognized within NASA and the industry that a clean-sheet vehicle design is required to demonstrate the theoretically predicted far-field aeroacoustic benefits of QuietSpike type morphing and other boom- mitigating concepts. NASA's Aeronautics Research Mission Directorate (ARMD) Supersonics Division has placed increased priority on near-term development and flight-testing of such a vehicle. To help achieve this objective, static and dynamic aerostructural proof-of-concept testing was considered a prudent step prior to a clean-sheet effort in order to reduce risk associated with a follow-on test program. Following a survey of potential test platforms, NASA Dryden's F-15B was selected as the target test vehicle primarily because of its unique ability to carry a largescale test apparatus to relevant supersonic flight speeds, so called the F15 -QS. The QuietSpike test article was constructed primarily of composite materials and attached to the forward fuselage of the F-1 5B bulkhead (see Figures 1,2). The QuietSpike test article replaces the current flight test noseboom and radome assembly. Power is supplied to the Quiet Spike motor assembly in order to extend and retract the Spike, and the Quiet Spike test article was appropriately instrumented with accelerometers, strain gages, pressure transducers, and thermocouples.

Description: The Aerostructures Test Wing (ATW) was developed to test unique concepts for flutter prediction and control synthesis. A follow-on to the successful ATW, denoted ATW2, was fabricated as a test bed to validate a variety of instrumentation in flight and to collect data for development of advanced signal processing algorithms for flutter prediction and aviation safety. As a means to estimate flutter speed, a ground vibration test (GVT) was performed. The results of a GVT are typically utilized to update structural dynamics finite element (FE) models used for flutter analysis. In this study, two GVT methodologies were explored to determine which nodes provide the best sensor locations: (i) effective independence and (ii) kinetic energy sorting algorithms. For measurement, ten and twenty sensors were used for three and 10 target test modes. A total of six accelerometer configurations measured frequencies and mode shapes. This included locations used in the original ATW GVT. Moreover, an optical measurement system was used to acquire data without mass effects added by conventional sensors. A considerable frequency shift was observed in comparing the data from the accelerometers to the optical data. The optical data provided robust data for use of the ATW2 finite element model update.

Description: As technologies advance, their growing complexity makes them harder to maintain. Detection methods for isolating and identifying impending problems are needed to balance this complexity. Through comparison of signal pairs from onboard sensors, the Beacon-based Exception Analysis For Multimissions (BEAM) algorithm can identify and help classify deviations in system operation from a data-trained statistical model. The goal of this task is to mature BEAM and validate its performance on a flying test bed. A series of F-18 flight demonstrations with BEAM monitoring engine parameters in real time was used to demonstrate in-the-field readiness. Captured F-18 and simulated F-18 engine data were used in model creation and training. The algorithm was then ported to the embedded system with a data buffering, file writing, and data-time-stamp monitoring shell to reduce the impact of embedded system faults on BEAM'S ability to correctly identify engine faults. Embedded system testing identified hardware related restrictions and contributed to iterative improvements in the code's runtime performance. The system was flown with forced engine flameouts and other pilot induced faults to simulate operation out of the norm. Successful detection of these faults, confirmed through post-flight data analysis, helped BEAM achieve TRL6.

Description: This DVD has several short videos showing some of the work that Dryden is involved in with experimental aircraft. These are: shots showing the Active AeroElastic Wing (AAW) loads calibration tests, AAW roll maneuvers, AAW flight control surface inputs, Helios flight, and takeoff, and Pathfinder takeoff, flight and landing.

Description: Measurement of average skin-friction coefficients have been made on six rocket-powered free-flight models by using the boundary-layer rake technique. The model configuration was the NACA RM-10, a 12.2-fineness-ratio parabolic body of revolution with a flat base. Measurements were made over a Mach number range from 1 to 3.7, a Reynolds number range 40 x 10(exp 6) to 170 x 10(exp 6) based on length to the measurement station, and with aerodynamic heating conditions varying from strong skin heating to strong skin cooling. The measurements show the same trends over the test ranges as Van Driest's theory for turbulent boundary layer on a flat plate. The measured values are approximately 7 percent higher than the values of the flat-plate theory. A comparison which takes into account the differences in Reynolds number is made between the present results and skin-friction measurements obtained on NACA RM-10 scale models in the Langley 4- by 4-foot supersonic pressure tunnel, the Lewis 8- by 6-foot supersonic tunnel, and the Langley 9-inch supersonic tunnel. Good agreement is shown at all but the lowest tunnel Reynolds number conditions. A simple empirical equation is developed which represents the measurements over the range of the tests.

Description: The author states some tentative principles in the absence of an existence theorem of sufficiently general solutions of viscous fluid equations which could be applied by engineers. These principles are used for the characterization of the singular points of flow, which can be determined and identified by experimental engineers to solve a problem of fluid mechanics by a model.

Description: This paper describes the experimental results concerning the detection of a crack in a rotating disk. The goal was to utilize blade tip clearance and shaft vibration measurements to monitor changes in the system's center of mass and/or blade deformation behaviors. The concept of the approach is based on the fact that the development of a disk crack results in a distorted strain field within the component. As a result, a minute deformation in the disk's geometry as well as a change in the system's center of mass occurs. Here, a notch was used to simulate an actual crack. The vibration based experimental results failed to identify the existence of a notch when utilizing the approach described above, even with a rather large, circumferential notch (l.2 in.) located approximately mid-span on the disk (disk radius = 4.63 in. with notch at r = 2.12 in.). This was somewhat expected, since the finite element based results in Part 1 of this study predicted changes in blade tip clearance as well as center of mass shifts due to a notch to be less than 0.001 in. Therefore, the small changes incurred by the notch could not be differentiated from the mechanical and electrical noise of the rotor system. Although the crack detection technique of interest failed to identify the existence ofthe notch, the vibration data produced and captured here will be utilized in upcoming studies that will focus on different data mining techniques concerning damage detection in a disk.

Description: An aircraft systems analysis was conducted to evaluate the net benefits of advanced technologies on two conceptual civil transport rotorcraft, to quantify the potential of future civil rotorcraft to become operationally viable and economically competitive, with the ultimate goal of alleviating congestion in our airways, runways and terminals. These questions are three of many that must be resolved for the successful introduction of civil transport rotorcraft: 1) Can civil transport rotorcraft actually relieve current airport congestion and improve overall air traffic and passenger throughput at busy hub airports? What is that operational scenario? 2) Can advanced technology make future civil rotorcraft economically competitive in scheduled passenger transport? What are those enabling technologies? 3) What level of investment is necessary to mature the key enabling technologies? This study addresses the first two questions, and several others, by applying a systems analysis approach to a broad spectrum of potential advanced technologies at a conceptual level of design. The method was to identify those advanced technologies that showed the most promise and to quantify their benefits to the design, development, production, and operation of future civil rotorcraft. Adjustments are made to sizing data by subject matter experts to reflect the introduction of new technologies that offer improved performance, reduced weight, reduced maintenance, or reduced cost. This study used projected benefits from new, advanced technologies, generally based on research results, analysis, or small-scale test data. The technologies are identified, categorized and quantified in the report. The net benefit of selected advanced technologies is quantified for two civil transport rotorcraft concepts, a Single Main Rotor Compound (SMRC) helicopter designed for 250 ktas cruise airspeed and a Civil Tilt Rotor (CTR) designed for 350 ktas cruise airspeed. A baseline design of each concept was sized for a representative civil passenger transport mission, using current technology. Individual advanced technologies are quantified and applied to resize the aircraft, thereby quantifying the net benefit of that technology to the rotorcraft. Estimates of development cost, production cost and operating and support costs are made with a commercial cost estimating program, calibrated to Boeing products with adjustments for future civil production processes. A cost metric of cash direct operating cost per available seat-mile (DOC ASM) is used to compare the cost benefit of the technologies. The same metric is used to compare results with turboprop operating costs. Reduced engine SFC was the most advantageous advanced technology for both rotorcraft concepts. Structural weight reduction was the second most beneficial technology, followed by advanced drive systems and then by technology for rotorcraft performance. Most of the technologies evaluated in this report should apply similarly to conventional helicopters. The implicit assumption is that resources will become available to mature the technologies for fullscale production aircraft. That assumption is certainly the weak link in any forecast of future possibilities. The analysis serves the purpose of identifying which technologies offer the most potential benefit, and thus the ones that should receive the highest priority for continued development. This study directly addressed the following NASA Subsonic Rotary Wing (SRW) subtopics: SR W.4.8.I.J Establish capability for rotorcraft system analysis and SRW. 4.8.I.4 Conduct limited technology benefit assessment on baseline rotorcraft configurations.

Description: NorthWest Research Associates (NWRA) has developed an Inverse Model for inverting aircraft wake vortex data. The objective of the inverse modeling is to obtain estimates of the vortex circulation decay and crosswind vertical profiles, using time history measurements of the lateral and vertical position of aircraft vortices. The Inverse Model performs iterative forward model runs using estimates of vortex parameters, vertical crosswind profiles, and vortex circulation as a function of wake age. Iterations are performed until a user-defined criterion is satisfied. Outputs from an Inverse Model run are the best estimates of the time history of the vortex circulation derived from the observed data, the vertical crosswind profile, and several vortex parameters. The forward model, named SHRAPA, used in this inverse modeling is a modified version of the Shear-APA model, and it is described in Section 2 of this document. Details of the Inverse Model are presented in Section 3. The Inverse Model was applied to lidar-observed vortex data at three airports: FAA acquired data from San Francisco International Airport (SFO) and Denver International Airport (DEN), and NASA acquired data from Memphis International Airport (MEM). The results are compared with observed data. This Inverse Model validation is documented in Section 4. A summary is given in Section 5. A user's guide for the inverse wake vortex model is presented in a separate NorthWest Research Associates technical report (Lai and Delisi, 2007a).

Description: One of the operational modes of the Terminal Area Simulation System (TASS) model simulates the three-dimensional interaction of wake vortices within turbulent domains in the presence of thermal stratification. The model allows the investigation of turbulence and stratification on vortex transport and decay. The model simulations for this work all assumed fully-periodic boundary conditions to remove the effects from any surface interaction. During the Base Period of this contract, NWRA completed generation of these datasets but only presented analysis for the neutral stratification runs of that set (Task 3.4.1). Phase 1 work began with the analysis of the remaining stratification datasets, and in the analysis we discovered discrepancies with the vortex time to link predictions. This finding necessitated investigating the source of the anomaly, and we found a problem with the background turbulence. Using the most up to date version TASS with some important defect fixes, we regenerated a larger turbulence domain, and verified the vortex time to link with a few cases before proceeding to regenerate the entire 25 case set (Task 3.4.2). The effort of Phase 2 (Task 3.4.3) concentrated on analysis of several scenarios investigating the effects of closely spaced aircraft. The objective was to quantify the minimum aircraft separations necessary to avoid vortex interactions between neighboring aircraft. The results consist of spreadsheets of wake data and presentation figures prepared for NASA technical exchanges. For these formation cases, NASA carried out the actual TASS simulations and NWRA performed the analysis of the results by making animations, line plots, and other presentation figures. This report contains the description of the work performed during this final phase of the contract, the analysis procedures adopted, and sample plots of the results from the analysis performed.

Description: This slide presentation reviews the flight test from the autonomous airborne refueling system. It includes information on the prototype system that can autonomously perform fueling, including during a turn or mild turbulence, and the autonomous rendezvous capability.

Description: A tank investigation has been conducted on a 1/8-size powered dynamic model of the Grumman JRF-5 airplane equipped with twin hydro-skis. The results of tests using two types of skis are presented: one had vertical sides joining the top surface to the chine; the other had the top surface faired to the chine to eliminate the vertical sides. Both configurations had satisfactory longitudinal stability although the model had a slightly greater stable elevator range available when the skis without the vertical sides were attached. Free model tests indicated no instability present when one ski emerged before the other. Considerable excess thrust was available at all speeds with either type of skis. A hump gross load-resistance ratio of 3.37 was obtained with the skis with the vertical sides and 3.53 with the other skis. Landing behavior in smooth water with yaw up to 15deg and roll up to 15deg in opposite directions was satisfactory with either type of skis.

Description: An investigation was made to determine the static longitudinal and lateral stability and control characteristics of a l/6-scale model of the revised Republic XF-84H airplane with and without the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and was equipped with a thin, three-blade supersonic-type propeller. Modifications incorporated in the revised model included a raised horizontal tail, increased rudder size, wing fences at 65 percent semispan, and a modified wing leading edge outboard of the fences. The test results for flap-retracted and flap-deflected conditions indicated that the revised configuration should be satisfactory for most normal flight conditions provided the angle of attack does not exceed the angle for pitch-up. An abrupt pitch-up tendency of the model was evident for the zero thrust condition above approximately 15' angle of attack. Although the effects of power were destabilizing, power-on longitudinal stability was satisfactory through the angle-of-attack range for which the model was stable with zero thrust. Above the angle of attack for pitch-up, an uncontrollable left roll-off tendency would be expected with power on and slats retracted. Projection of wing slats or use of leading-edge chord-extensions with only the left extension drooped were found beneficial in controlling the roll-off tendency with power on; however the most effective means found was projection of only the left slat.

Description: The Ko displacement theory, formulated for weak nonuniform (slowly changing cross sections) cantilever beams, was applied to the deformed shape analysis of the doubly-tapered wings of the Ikhana unmanned aircraft. The two-line strain-sensing system (along the wingspan) was used for sensing the bending strains needed for the wing-deformed shapes (deflections and cross-sectional twist) analysis. The deflection equation for each strain-sensing line was expressed in terms of the bending strains evaluated at multiple numbers of strain-sensing stations equally spaced along the strain-sensing line. For the preflight shape analysis of the Ikhana wing, the strain data needed for input to the displacement equations for the shape analysis were obtained from the nodal-stress output of the finite-element analysis. The wing deflections and cross-sectional twist angles calculated from the displacement equations were then compared with those computed from the finite-element computer program. The Ko displacement theory formulated for weak nonlinear cantilever beams was found to be highly accurate in the deformed shape predictions of the doubly-tapered Ikhana wing.

Description: Free-flight tests have been made to determine the zero-lift drag of several configurations of the XAAM-N-2 pilotless aircraft. Base-pressure measurements were also obtained for some of the configurations. The results show that increasing the wing-thickness ratio from 4 to 6 percent increased the wing drag by about 100 percent at M = 1.3 and by about 30 percent at M = 1.8. Increasing the nose fineness ratio from 5.00 to 6.25 reduced the drag coefficient of the wingless models a maximum of about 0.030 (10 percent) at M = 2.0. A corresponding change in nose shape for the winged models decreased the drag coefficient by about 0.05 in the Mach number range from 1.1 to 1.4; at Mach numbers greater than 1.6 no measurable reduction in drag coefficient was obtained. The drag of the present Sparrow fuselage is less than that of a parabolic fuselage which could contain the same equipment.

Description: A limited investigation of a 1/24-scale dynamically similar model of the Navy Bureau of Aeronautics DR-77 design was conducted in Langley tank no. 2 to determine the calm-water take-off and the rough-water landing characteristics of the design with particular regard to the take-off resistance and the landing accelerations. During the take-off tests, resistance, trim, and rise were measured and photographs were taken to study spray. During the landing tests, motion-picture records and normal-acceleration records were obtained. A ratio of gross load to maximum resistance of 3.2 was obtained with a 30 deg. dead-rise hydro-ski installation. The maximum normal accelerations obtained with a 30 deg. dead-rise hydro-ski installation were of the order of 8g to log in waves 8 feet high (full scale). A yawing instability that occurred just prior to hydro-ski emergence was improved by adding an afterbody extension, but adding the extension reduced the ratio of gross load to maximum resistance to 2.9.

Description: The stator-blade angles in the twelfth through fifteenth stages of a 16-stage axial-flow compressor were increased 3O. The over-all performance of this modified compressor is compared to the performance of the compressor with original blade angles. The matching characteristics of the modified compressor and a two-stage turbine were obtained and compared to those of the compressor with original blade angles and the same turbine.

Description: An investigation has been conducted to determine the static stability and control and damping in roll and yaw of a 0.13-scale model of the Convair XFY-1 airplane with propellers off from 0 deg to 90 deg angle of attack. The tests showed that a slightly unstable pitch-up tendency occurred simultaneously with a break in the normal-force curve in the angle-of-attack range from about 27 deg to 36 deg. The top vertical tail contributed positive values of static directional stability and effective dihedral up to an angle of attack of about 35 deg. The bottom tail contributed positive values of static directional stability but negative values of effective dihedral throughout the angle-of-attack range. Effectiveness of the control surfaces decreased to very low values at the high angles of attack, The model had positive damping in yaw and damping in roll about the body axes over the angle-of-attack range but the damping in yaw decreased to about zero at 90 deg angle of attack.

Description: Altitude performance characteristics of the J65-B3 turbojet engine and its components were obtained at engine-inlet conditions corresponding to Reynolds number indices from 0.2 to 0.8 over a range of corrected engine speeds from 70 to 110 percent of rated speed. Engine operational limits up to an altitude of 75,000 feet together with ignition and windmilling characteristics were also obtained. The engine and component data are presented both in graphical and in tabulated form. The operational characteristics are presented in graphical form.

Description: A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion system and compared the results to those of the baseline single thruster system.

Description: A series of tests were conducted on the electromechanical actuators of the X-43A research vehicle in preparation for the Mach 7 and 10 hypersonic flights. The tests were required to help validate the actuator models in the simulation and acquire a better understanding of the installed system characteristics. Static and dynamic threshold, multichannel crosstalk, command-to-surface timing, free play, voltage regeneration, calibration, frequency response, compliance, hysteretic damping, and aircraft-in-the-loop tests were performed as part of this effort. This report describes the objectives, configurations, and methods for those tests, as well as the techniques used for developing second-order actuator models from the test results. When the first flight attempt failed because of actuator problems with the launch vehicle, further analysis and model enhancements were performed as part of the return-to-flight activities. High-fidelity models are described, along with the modifications that were required to match measurements taken from the research vehicle. Problems involving the implementation of these models into the X-43A simulation are also discussed. This report emphasizes lessons learned from the actuator testing, simulation modeling, and integration efforts for the X-43A hypersonic research vehicle.

Description: Independently deflectable control surfaces are located on the trailing edge of the wing of a blended wing-body aircraft. The reconfiguration control system of the present invention controls the deflection of each control surface to optimize the spanwise lift distribution across the wing for each of several flight conditions, e.g., cruise, pitch maneuver, and high lift at low speed. The control surfaces are deflected and reconfigured to their predetermined optimal positions when the aircraft is in each of the aforementioned flight conditions. With respect to cruise, the reconfiguration control system will maximize the lift to drag ratio and keep the aircraft trimmed at a stable angle of attack. In a pitch maneuver, the control surfaces are deflected to pitch the aircraft and increase lift. Moreover, this increased lift has its spanwise center of pressure shifted inboard relative to its location for cruise. This inboard shifting reduces the increased bending moment about the aircraft's x-axis occasioned by the increased pitch force acting normal to the wing. To optimize high lift at low speed, during take-off and landing for example, the control surfaces are reconfigured to increase the local maximum coefficient of lift at stall-critical spanwise locations while providing pitch trim with control surfaces that are not stall critical.

Description: A simulation of a commercial engine has been developed in a graphical environment to meet the increasing need across the controls and health management community for a common research and development platform. This paper describes the Commercial Modular Aero Propulsion System Simulation (C-MAPSS), which is representative of a 90,000-lb thrust class two spool, high bypass ratio commercial turbofan engine. A control law resembling the state-of-the-art on board modern aircraft engines is included, consisting of a fan-speed control loop supplemented by relevant engine limit protection regulator loops. The objective of this paper is to provide a top-down overview of the complete engine simulation package.

Description: This paper has summarized recent NASA research into scaling of SLD conditions with data from both SLD and Appendix C tests. Scaling results obtained by applying existing scaling methods for size and test-condition scaling will be reviewed. Large feather growth issues, including scaling approaches, will be discussed briefly. The material included applies only to unprotected, unswept geometries. Within the limits of the conditions tested to date, the results show that the similarity parameters needed for Appendix C scaling also can be used for SLD scaling, and no additional parameters are required. These results were based on visual comparisons of reference and scale ice shapes. Nearly all of the experimental results presented have been obtained in sea-level tunnels. The currently recommended methods to scale model size, icing limit and test conditions are described.

Description: The National Aeronautics and Space Administration Dryden Flight Research Center has a long history in developing simulations of experimental fixed-wing aircraft from gliders to suborbital vehicles on platforms ranging from desktop simulators to pilot-in-the-loop/aircraft-in-the-loop simulators. Regardless of the aircraft or simulator hardware, much of the software framework is common to all NASA Dryden simulators. Some of this software has withstood the test of time, but in recent years the push toward high-fidelity user-friendly simulations has resulted in some significant changes. This report presents an overview of the current NASA Dryden simulation software framework and capabilities with an emphasis on the new features that have permitted NASA to develop more capable simulations while maintaining the same staffing levels.

Description: The Aviation System Monitoring and Modeling (ASMM) Project was one of the projects within NASA s Aviation Safety Program from 1999 through 2005. The objective of the ASMM Project was to develop the technologies to enable the aviation industry to undertake a proactive approach to the management of its system-wide safety risks. The ASMM Project entailed four interdependent elements: (1) Data Analysis Tools Development - develop tools to convert numerical and textual data into information; (2) Intramural Monitoring - test and evaluate the data analysis tools in operational environments; (3) Extramural Monitoring - gain insight into the aviation system performance by surveying its front-line operators; and (4) Modeling and Simulations - provide reliable predictions of the system-wide hazards, their causal factors, and their operational risks that may result from the introduction of new technologies, new procedures, or new operational concepts. This report is a documentation of the history of this highly successful project and of its many accomplishments and contributions to improved safety of the aviation system.

Description: A high bypass jet engine fan case represents one of the largest, heaviest single components in an engine. In addition to supporting the inlet and providing the fan flowpath, the most critical function is the containment of a failed fan blade. In this development program, a lightweight, low-cost composite containment case with diagnostic capabilities was developed, fabricated, and tested. The fan case design, containment methods, and diagnostic concepts evaluated in the initial Propulsion 21 program were improved and scaled up to a full case design.

Description: As gas foil journal bearings become more prevalent in production machines, such as small gas turbine propulsion systems and microturbines, system-level performance issues must be identified and quantified in order to provide for successful design practices. Several examples of system-level design parameters that are not fully understood in foil bearing systems are thermal management schemes, alignment requirements, balance requirements, thrust load balancing, and others. In order to address some of these deficiencies and begin to develop guidelines, this paper presents a preliminary experimental investigation of the misalignment tolerance of gas foil journal bearing systems. Using a notional gas foil bearing supported rotor and a laser-based shaft alignment system, increasing levels of misalignment are imparted to the bearing supports while monitoring temperature at the bearing edges. The amount of misalignment that induces bearing failure is identified and compared to other conventional bearing types such as cylindrical roller bearings and angular contact ball bearings. Additionally, the dynamic response of the rotor indicates that the gas foil bearing force coefficients may be affected by misalignment.

Description: A critical aspect of the Access 5 program is identifying appropriate spectrum for civil and commercial purposes. However, currently, there is no spectrum allocated for the command/control link between the aircraft control station and the unmanned aircraft. Until such frequency spectrum is allocated and approved, it will be difficult for the UAS community to obtain civil airworthiness certification and operate in the NAS on a routine basis. This document provides a perspective from the UAS community on Agenda Items being considered for the upcoming World Radiocommunication Conference 2007 (WRC 07). Primarily, it supports the proposal to add Aeronautical Mobile (Route) Services (AM(R)S) to existing bands that could be used for UAS Line-of-Sight operations. It also recommends the need to identify spectrum that could be used for an Aeronautical Mobile Satellite (Route) Service (AMS(R)S) that would allow UAS to operate Beyond Line-of-Sight. If spectrum is made available to provide these services, it will then be incumbent upon the UAS community to justify their use of this spectrum as well as the assurance that they will not interfere with other users of this newly allocated spectrum.

Description: This document provides a study of the technical literature related to Command and Control (C2) link security for Unmanned Aircraft Systems (UAS) for operation in the National Airspace System (NAS). Included is a preliminary set of functional requirements for C2 link security.

Description: This document describes a method to demonstrate that a UAS, operating in the NAS, can avoid collisions with an equivalent level of safety compared to a manned aircraft. The method is based on the calculation of a collision probability for a UAS , the calculation of a collision probability for a base line manned aircraft, and the calculation of a risk ratio given by: Risk Ratio = P(collision_UAS)/P(collision_manned). A UAS will achieve an equivalent level of safety for collision risk if the Risk Ratio is less than or equal to one. Calculation of the probability of collision for UAS and manned aircraft is accomplished through event/fault trees.

Description: The National Aeronautics and Space Administration (NASA) Access 5 Project Office sponsored a cooperative collision avoidance flight demonstration program for unmanned aircraft systems (UAS). This flight test was accomplished between September 21st and September 27th 2005 from the Mojave Airport, Mojave, California. The objective of these flights was to collect data for the Access 5 Cooperative Collision Avoidance (CCA) Work Package simulation effort, i.e., to gather data under select conditions to allow validation of the CCA simulation. Subsequent simulation to be verified were: Demonstrate the ability to detect cooperative traffic and provide situational awareness to the ROA pilot; Demonstrate the ability to track the detected cooperative traffic and provide position information to the ROA pilot; Demonstrate the ability to determine collision potential with detected cooperative traffic and provide notification to the ROA pilot; Demonstrate that the CCA subsystem provides information in sufficient time for the ROA pilot to initiate an evasive maneuver to avoid collision; Demonstrate an evasive maneuver that avoids collision with the threat aircraft; and lastly, Demonstrate the ability to assess the adequacy of the maneuver and determine that the collision potential has been avoided. The Scaled Composites, LLC Proteus Optionally Piloted Vehicle (OPV) was chosen as the test platform. Proteus was manned by two on-board pilots but was also capable of being controlled from an Air Vehicle Control Station (AVCS) located on the ground. For this demonstration, Proteus was equipped with cooperative collision sensors and the required hardware and software to place the data on the downlink. Prior to the flight phase, a detailed set of flight test scenarios were developed to address the flight test objectives. Two cooperative collision avoidance sensors were utilized for detecting aircraft in the evaluation: Traffic Alert and Collision Avoidance System-II (TCAS-II) and Automatic Dependent Surveillance Broadcast (ADS-B). A single intruder aircraft was used during all the flight testing, a NASA Gulfstream III (G-III). During the course of the testing, six geometrically different near-collision scenarios were evaluated. These six scenarios were each tested using various combinations of sensors and collision avoidance software. Of the 54 planned test points 49 were accomplished successfully. Proteus flew a total of 21.5 hours during the testing and the G-III flew 19.8 hours. The testing fully achieved all flight test objectives. The Flight IPT performed an analysis to determine the accuracy of the simulation model used to predict the location of the host aircraft downstream during an avoidance maneuver. The data collected by this flight program was delivered to the Access 5 Cooperative Collision Avoidance (CCA) Work Package Team who was responsible for reporting on their analysis of this flight data.

Description: The present invention discloses systems and methods for the performance enhancement of rotary wing aircraft through reduced torque, noise and vibration. In one embodiment, a system includes a sail having an aerodynamic shape positioned proximate to a tip of the rotor blade. An actuator may be configured to rotate the sail relative to the blade tip. a A control system receives information from a rotorcraft system and commands the actuator to rotate the sail to a predetermined favorable rotor blade operating condition. In another embodiment, a method includes configuring the rotorcraft in a selected flight condition, communicating input signals to a control system operable to position sails coupled to tips of blades of a rotor assembly, processing the input signals according to a constraint condition to generate sail positional information, and transferring the sail positional information to the sail.