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An integration study was performed coupling an SP-100 reactor with either a Brayton or Stirling power conversion subsystem. A power level of 100 kWe was selected for the study. The power system was to be compatible with both the lunar and Mars surface environment and require no site preparation. In addition, the reactor was to have integral shielding and be completely self-contained, including its own auxiliary power for start-up. Initial reliability studies were performed to determine power conversion redundancy and engine module size. Previous studies were used to select the power conversion optimum operating conditions (ratio of hot-side temperature to cold-side temperature). Results of the study indicated that either the Brayton or Stirling power conversion subsystems could be integrated with the SP-100 reactor for either a lunar or Mars surface power application.

A first attempt to study civil aircraft operations comprehensively, prior to having the airplane, occurred before the initial operation of U.S. subsonic jets. One airline carried out a manual-simulated “paper jet” operation lasting fifteen months. Today, computerized simulation of machines, methods, and operations has become commonplace, and replaces the slide rule and tedious day-by-day inputs of aircraft operational criteria. Computerized simulations are also applied to every aspect of the SST design and operations. These are important, but the results being should be used with caution and judgement.

Governmental assurance documentation bibliography updated; new tabulation effective as of April 1, 1967. Latest revision indicated in all instances, but no attempt was made to list supplements or amendments. Department of Defense Index of Specifications and Standards (DODISS) published annually in three parts (alphabetic, numerical, and listing of Federal Supply Classification following unclassified documents.

The Cardinal is a Super Short Takeoff and Landing (SSTOL) aircraft, which is designed to fulfill the desire for center-city to center-city travel by utilizing river “barges” for short takeoffs and landings to avoid construction of new runways or heliports. In addition, the Cardinal will fulfill the needs of the U.S. Navy for a Carrier On-board Delivery (COD) aircraft to replace the C-2 Greyhound. Design requirements for the Cardinal included a takeoff ground roll of 300 ft, a landing ground roll of 400 ft, cruise at 350 knots with a range of up to 1500 nm with reserves, payload of 24 passengers and baggage for a commercial version or a military version with a 10,000 lb payload, capable of carrying two GE F110 engines for the F-14D, and a spot factor requirement of 60 feet by 29 feet.

Most of current jet aircraft circulate fuel on the airframe to match heat loads with available heat sink. The demands for thermal management in wide range of air vehicle systems are growing rapidly along with the increased mission power, vehicle survivability, flight speeds, and so on. With improved aircraft performance and growth of heat load created by Aircraft Mounted Accessory Drive (AMAD) system and hydraulic system, effectively removing the large amount of heat load on the aircraft is gaining crucial importance. Fuel is becoming heat transfer fluid of choice for aircraft thermal management since it offers improved heat transfer characteristics and offers fewer system penalties than air. In the scope of this paper, an AMESim model is built which includes airframe fuel and hydraulic systems with AMAD gearbox of a jet trainer aircraft. The integrated model will be evaluated for thermal performance.

The paper describes a numerical study, supported by experiments, on light aircraft 2-Stroke Direct Injected Diesel engines, typically rated up to 110 kW (corresponding to about 150 imperial HP). The engines must be as light as possible and they are to be directly coupled to the propeller, without reduction drive. The ensuing main design constraints are: i) in-cylinder peak pressure as low as possible (typically, no more than 120 bar); ii) maximum rotational speed limited to 2600 rpm. As far as exhaust emissions are concerned, piston aircraft engines remain unregulated but lack of visible smoke is a customer requirement, so that a value of 1 is assumed as maximum Smoke number. For the reasons clarified in the paper, only three cylinder in line engines are investigated. Reference is made to two types of scavenging and combustion systems, designed by the authors with the assistance of state-of-the-art CFD tools and described in detail in a parallel paper.

Aircraft potable (drinking) water systems haven’t changed significantly in the last half-century. These systems consist of cylindrical water tanks pressurized by bleed air from the jet engines, with insulated stainless steel distribution lines. What has changed recently is the increase in the possibility of aircraft picking up contaminated drinking water at foreign and domestic stops. Customer awareness of these problems has also changed - to the point where having reliable drinking water is now a competitive issue among airlines. Old style potable water systems that are used on modern aircraft are high maintenance and exacerbate the growth of microbes because the water is static much of the time. The integrity of some pressurized water tanks are also a concern after years of use. Cost-effective mechanical and biological solutions exist that can significantly reduce the amount of chemicals added and provide good potable water.

Preliminary results from testing of 26 X 6.6 radial-belted and bias-ply aircraft tires at NASA Langley's Aircraft Landing Dynamics Facility (ALDF) are reviewed. These tire tests are part of a larger, on going joint NASA/FAA/Industry Surface Traction and Radial Tire (START) Program involving three different tire sizes. The 26 X 6.6 tire size evaluation includes cornering performance tests throughout the aircraft ground operational speed range for both dry and wet runway surfaces. Static test results to define 26 X 6.6 tire vertical stiffness properties are also presented and discussed.

Sundstrand is investigating 270-Vdc/hybrid 115-Vac electrical power generating and distribution systems technology so as to be well prepared to offer such systems for future aircraft applications. The approach taken has been to design, build, and test a representative system that meets or exceeds the tightest of the performance standards as defined by miliary standards. This paper describes a single-channel, 120-kW hybrid system and presents some typical performance data. The dc bus supplies a 30-kVA, 400-Hz, 115-Vac inverter; constant power load banks of up to 150 kW; and a resistive load bank of up to 90 kW. System simulation studies indicated the potential for unstable operation due to the negative impedance of the constant power load in conjunction with the source ripple filter and the load EMI filters. Unstable voltage and current were observed in system testing when the magnitude of the source impedance was not sufficiently below that of the load impedance.

Sundstrand has been investigating 270-Vdc/hybrid 115-Vac electrical power generating systems (EPGS) technology in preparation for meeting the electrical power generating system (EPGS) requirements for future aircraft (1). Systems such as the one being investigated are likely to be suitable for the More-Electric Aircraft (MEA) concepts presently under industry and military study. The present Sundstrand single-channel testbed is being further expanded to better understand the electrical system performance characteristics and power quality requirements of an MEA in which traditional mechanical subsystems are replaced by those of a “more-electric” nature. This paper presents the most recent Sundstrand 270-Vdc system transient performance data, and describes the modifications being made to the 270-Vdc/hybrid 115-Vac testbed.

Different Airbus Helicopters main rotor blade profiles were tested in different icing wind tunnels and for different icing conditions. One of the objectives of the accretion tests was to validate the use of 2D icing scaling laws established for fixed wing aircraft on helicopter blade profiles. Therefore, ice shapes resulting from tests with the same icing similarity parameters are compared to each other allowing the assessment of icing scaling laws for helicopter applications. This paper presents the icing scaling laws used at Airbus Helicopters on blade profiles, the different test set ups and test models and it presents the comparison of the ice shapes collected during the icing wind tunnel test campaigns.

A helicopter blade profile was tested in the DGA Aero-engine Testing's icing altitude test facility S1 in Saclay, France during the winter of 2013/2014. The airfoil was a helicopter main rotor OA312 blade profile made out of composite material and with a metallic erosion shield. Dry air and ice accretion tests have been performed in order to assess the iced airfoil's aerodynamic behaviour. Several icing conditions were tested up through Mach numbers around 0.6. This paper presents the test setup, the test model and some of the test results. The test results presented in this paper include the ice shapes generated as well as dry air and iced airfoil lift and drag curves (polars) which were obtained with the real ice shapes on the airfoil.

The term “3 inch ice shapes” has assumed numerous definitions throughout the years. At times it has been used to generally characterize large glaze ice accretions on the major aerodynamic surfaces (wing, horizontal stabilizer, vertical stabilizer) for evaluating aerodynamic performance and handling qualities after a prolonged icing encounter. It has also been used as a more direct criterion while determining or enforcing sectional ice shape characteristics such as the maximum pinnacle height. It is the authors’ observation that over the years, the interpretation and application of this term has evolved and is now broadly misunderstood. Compounding the situation is, at present, a seemingly contradictory set of guidance among (and even within) the various international regulatory agencies resulting in an ambiguous set of expectations for design and certification specialists.

Accurate measurement of countersinks in curved parts has always been a challenge. The countersink reference is defined relative to the panel surface which includes some degree of curvature. This curvature thus makes accurate measurements very difficult using both contact and 2D non-contact measurements. By utilizing structured light 3D vision technologies, the ability to very accurately measure a countersink to small tolerances can be achieved. By knowing the pose of the camera and projector, triangulation can be used to calculate the distance to thousands of points on the panel and countersink surface. The plane of the panel is then calculated using Random Sample Consensus (RANSAC) method from the dataset of points which can be adjusted to account for panel curvatures. The countersink is then found using a similar RANSAC method.

The objective of this project was to replace electromechanical power line contactors with a Static Transfer Switch (STS) to improve the transfer of electrical power between aircraft generators and decrease required maintenance. The switch requirements include high reliability, lightweight, and high speed (less than 15mS) power transfer. An STS can shorten the bus transfer time to less than the “ride-through” of aircraft electronic loads and therefore have the ability to control and transfer electrical power while maintaining critical mission requirements. The content of this paper and presentation will discuss the initial problem, the research and development approach, design, and initial testing of the STS.

Vertical-Junction-Field-Effect-Transistors (VJFETs) are currently the most mature SiC devices for high power/temperature switching. High-voltage VJFETs are typically designed normally-on to ensure voltage control operation at high current-gain. However, to exploit the high voltage/temperature capabilities of VJFETs in a normally-off high-current voltage-controlled switch, high-voltage normally-on and low-voltage normally-off VJFETs were connected in the cascode configuration. In this paper, we review the high temperature DC characteristics of VJFETs and 1200 V normally-off cascode switches. The measured parameter shifts in the 25°C to 300°C temperature range are in excellent agreement with theory, confirming fabrication of robust SiC VJFETs and cascode switches.

Previous Flex Track drilling systems move along two parallel tracks that conform to the contour of a work piece surface. Until recently, applications have been limited to relatively simple surfaces such as the cylindrical mid-body fuselage join of a commercial aircraft. Recent developments in the state of the art have introduced the 5-axis variant which is capable of precision drilling on complex contours. This paper presents solutions to two positioning challenges associated with this added functionality: the ability to align the spindle axis normal to an angled drilling surface while maintaining accuracy in tool-point position, the ability to maintain synced motion between dual drives on complex track profiles.

The transition towards More Electric Aircraft (MEA) architectures has challenges relating to integration of power electronics with the starter generator system for on-engine application. To efficiently operate the power electronics in the hostile engine environment at high switching frequency and for better thermal management, use of silicon carbide (SiC) power devices for a bi-directional power converter is examined. In this paper, development of a 50 kVA bi-directional converter operating at an ambient temperature of about 2000C is presented. The design and operation of the converter with details of control algorithm implementation and cooling chamber design are also discussed.

As a part of a program jointly supported by the USAF and Canada's Department of National Defense, BlueStar is developing large (50 to 100-Ah)lithium ion cells for aircraft and spacecraft applications. Presently, 20-Ah cells are being developed as the first stage of the scale-up process and the design of these cells involves several tradeoffs related to the specific nature of this application. This paper will present the design of this first generation cell in the context of these tradeoffs as well as presenting the results of the performance and life testing of these cells.

Enemy threats during Operation Desert Storm drove many Allied ground attack aircraft to medium altitude to deliver their weapons. Although many Precision Guided Munitions (PGMs) proved to be highly accurate at increased heights, the majority of bombs dropped from medium altitude were low-drag general purpose bombs (LDGP). The aircraft carrying the LDGPs traded accuracy for altitude in order to reduce the risk to aircraft and crews. This event helped highlight the need to drop LDGP weapons more accurately. Additionally, there are environments where even many current PGMs are ineffective. During the Gulf War, for example, smoke from the Kuwaiti oil fires obscured the skies, prohibiting the use of laser guided bombs. Currently there is a program in testing that is designed to fill these operational gaps. The Joint Direct Attack Munition (JDAM) is being developed to convert our inventory of LDGPs into PGMs and modeling/simulation is playing a vital role in its design.