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Helicopter external air transportation plays an important role in today's world. For both military and civilian helicopters, external sling load operations offer an efficient and expedient method of handling heavy, oversized cargo. With the ability to reach areas otherwise inaccessible by ground transportation, helicopter external load operations are conducted in industries such as logging, construction, and fire fighting, as well as in support of military tactical transport missions. Historically, helicopter and load combinations have been qualified through flight testing, requiring considerable time and cost. With advancements in simulation and flight test techniques there is potential to substantially reduce costs and increase the safety of helicopter sling load certification. Validated simulation tools make possible accurate prediction of operational flight characteristics before initial flight tests. Real time analysis of test data improves the safety and efficiency of the testing programs. To advance these concepts, the U.S. Army and NASA, in cooperation with the Israeli Air Force and Technion, under a Memorandum of Agreement, seek to develop and validate a numerical model of the UH-60 with sling load and demonstrate a method of near real time flight test analysis. This thesis presents results from flight tests of a U.S. Army Black Hawk helicopter with various external loads. Tests were conducted as the U.S. first phase of this MOA task. The primary load was a container express box (CONEX) which contained a compact instrumentation package. The flights covered the airspeed range from hover to 70 knots. Primary maneuvers were pitch and roll frequency sweeps, steps, and doublets. Results of the test determined the effect of the suspended load on both the aircraft's handling qualities and its control system's stability margins. Included were calculations of the stability characteristics of the load's pendular motion. Utilizing CIFER(R) software, a method for near-real time system identification was also demonstrated during the flight test program.
The initial flight test phase of the modified C-8A airplane was conducted. The primary objective of the testing was to establish the basic airworthiness of the research vehicle. This included verification of the structural design and evaluation of the aircraft's systems. Only a minimum amount of performance testing was scheduled; this has been used to provide a preliminary indication of the airplane's performance and flight characteristics for future flight planning. The testing included flutter and loads investigations up to the maximum design speed. The operational characteristics of all systems were assessed including hydraulics, environmental control system, air ducts, the vectoring conical nozzles, and the stability augmentation system (SAS). Approaches to stall were made at three primary flap settings: up, 30 deg and 65 deg, but full stalls were not scheduled. Minimum control speeds and maneuver margins were checked. All takeoffs and landings were conventional, and STOL performance was not scheduled during this phase of the evaluation.
The Space Shuttle orbiter flight test program has required the aerodynamicist to take a new approach in determining flight characteristics. A conventional flight test program of slowly and cautiously approaching more severe flight conditions was not possible with the orbiter. On the first four orbital flights, The orbiter entered the atmosphere at approximately Mach 29 and decelerated through the Mach range (the subsonic portion of flight had also been flown by another orbiter vehicle during the Approach and Landing Test Program). Certification for these flights was achieved by an extensive wind tunnel test and analysis program. The initial series of flights of the orbiter were heavily instrumented for the purpose of obtaining accurate aerodynamic data. The flight data derived from the entry Mach range provided comparisons between flight and wind tunnel derived predicted data in the areas of both aerodynamic performance and longitudinal trim.
An investigation was conducted to determine the low-speed flight dynamic behavior of a representative advanced turboprop business/commuter aircraft concept. Free-flight tests were conducted in the NASA Langley Research Center's 30- by 60-Foot Tunnel. In support of the free-flight tests, conventional static, dynamic, and free-to-roll oscillation tests were performed. Tests were intended to explore normal operating and post stall flight conditions, and conditions simulating the loss of power in one engine.
A flight test of a business jet aircraft with various ice accretions was performed to obtain data to validate flight simulation models developed through wind tunnel tests. Three types of ice accretions were tested: pre-activation roughness, runback shapes that form downstream of the thermal wing ice protection system, and a wing ice protection system failure shape. The high fidelity flight simulation models of this business jet aircraft were validated using a software tool called "Overdrive." Through comparisons of flight-extracted aerodynamic forces and moments to simulation-predicted forces and moments, the simulation models were successfully validated. Only minor adjustments in the simulation database were required to obtain adequate match, signifying the process used to develop the simulation models was successful. The simulation models were implemented in the NASA Ice Contamination Effects Flight Training Device (ICEFTD) to enable company pilots to evaluate flight characteristics of the simulation models. By and large, the pilots confirmed good similarities in the flight characteristics when compared to the real airplane. However, pilots noted pitch up tendencies at stall with the flaps extended that were not representative of the airplane and identified some differences in pilot forces. The elevator hinge moment model and implementation of the control forces on the ICEFTD were identified as a driver in the pitch ups and control force issues, and will be an area for future work.
The physical parameters of the flight simulator visual system that characterize the system and determine its fidelity are identified and defined. The characteristics of visual simulation systems are discussed in terms of the basic categories of spatial, energy, and temporal properties corresponding to the three fundamental quantities of length, mass, and time. Each of these parameters are further addressed in relation to its effect, its appropriate units or descriptors, methods of measurement, and its use or importance to image quality.
Previous work in the development of quantitative models for the prediction of passenger reaction to motion and vehicle environment parameters in flight was extended to include a class of aircraft appropriate for low-density, short-haul service. The results indicate that it is possible to obtain quantitative response inputs from an usually small special test-subject group which will be representative of the general traveling public. Additional data which indicate the importance of comfort as a factor in evaluating ride quality was obtained, and identification of the factors which contribute to judgments regarding comfort level was improved. Seat comfort and seat spacing is very vital in the smaller aircraft. Mathematical modeling applied in conjuction with passenger reaction data was shown to be very useful for establishing ride-quality design criteria.
In support of new human rated spacecraft development programs, the Mission Operations Directorate at NASA Johnson Space Center has implemented a formal method for the assessment of spacecraft operability. This "Spacecraft Flight Operability Assessment Scale" defines six key themes of flight operability, with guiding principles and goals stated for each factor. A standardized rating technique provides feedback that is useful to the operations, design and program management communities. Applicability of this concept across the program structure and life cycle is addressed. Examples of operationally desirable and undesirable spacecraft design characteristics are provided, as is a sample of the assessment scale product.
Flight characteristics, and spacecraft performance for Fire Flight II project
At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the Flight Dynamics Branch at the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted in the Langley 12 foot Low Speed Wind Tunnel to identify and improve the stability and control characteristics of the vehicle. The wind tunnel test resulted in several configuration modifications which included increased elevator size, increased vertical tail size and tail moment arm, increased rudder size and aileron size, the addition of vertical wing tip fins, and the addition of leading-edge droops on the outboard wing panel to improve stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.
At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted to identify and improve the stability and control characteristics of the vehicle. The wind-tunnel test resulted in several configuration modifications which included increased elevator area, increased vertical tail area and moment arm, increased rudder area and aileron area, the addition of vertical wing-tip fins, and the addition of leading-edge droops on the outboard wing panel to improve the stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.
The SUbsonic Single Aft eNgine (SUSAN) Electrofan is a novel aircraft concept which utilizes a single aft-mounted engine, electrified aircraft propulsion, an emergency backup battery, as well as state-of-the-art aerodynamic design and thermal management systems to reduce the overall environmental impact of the aircraft. Mission profiles, which define the aircraft state and flight characteristics throughout various phases of flight, are an important component of aircraft conceptual design. These defined characteristics can serve as sizing cases or design constraints. By defining the mission profiles to be within airworthiness certification standards and regulations, aircraft designers can ensure the conceptual design is predicted to be within compliance of these regulations throughout the design process. This paper will present mission profiles for the SUSAN Electrofan aircraft, including the basis for defining the characteristics required during each phase of flight.
The Subsonic Single Aft engine (SUSAN) Electrofan is a novel aircraft concept which utilizes a single aft-mounted engine, electrified aircraft propulsion, an emergency backup battery, as well as state-of-the-art aerodynamic design and thermal management systems to reduce the overall environmental impact of the aircraft. Mission profiles, which define the aircraft state and flight characteristics throughout various phases of flight, are an important component of aircraft conceptual design. These defined characteristics can serve as sizing cases or design constraints. By defining the mission profiles to be within airworthiness certification standards and regulations, aircraft designers can ensure the conceptual design is predicted to be within compliance of these regulations throughout the design process. This paper will present mission profiles for the SUSAN Electrofan aircraft, including the basis for defining the characteristics required during each phase of flight.
Mathematical models of human pilot performance in roll tracking task derived from flight simulator and T 33 aircraft situations
Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.
The characteristics of the Langley 34 cm (13.5 in.) pilot cryogenic transonic pressure tunnel are described, and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8,600,000 obtained first at a stagnation pressure of 4.91 atmospheres at a stagnation temperature of +120 F and then at a stagnation pressure of 1.19 atmospheres at a stagnation temperature of -250 F.
Basic concepts involved in the mathematical modeling of the aerodynamic response of an aircraft to arbitrary maneuvers are reviewed. The original formulation of an aerodynamic response in terms of nonlinear functionals is shown to be compatible with a derivation based on the use of nonlinear functional expansions. Extensions of the analysis through its natural connection with ideas from bifurcation theory are indicated.