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A 10 KWe dual-mode space power system concept has been identified which is based on INEL's Small Externally-fueled Heat Pipe Thermionic Reactor (SEHPTR) concept. This power system will enhance user capabilities by providing reliable electric power and by providing two propulsion systems; electric power for an arc-jet electric propulsion system and direct thrust by heating hydrogen propellant inside the reactor. The low thrust electric thrusters allow efficient station keeping and long-term maneuvering. The direct thrust capability can provide tens of pounds of thrust at a specific impulse of around 730 seconds for maneuvers that must be performed more rapidly. The direct thrust allows the nuclear power system to move a payload from Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO) in less than one month using approximately half the propellant of a cryogenic chemical stage.

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.

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.

A 500 hours endurance test was performed with a heavy-duty engine (Euro IV); MAN type D 0836 LFL 51 equipped with a PM-Kat®. As fuel 100% biodiesel was used that met the European specification EN 14214. The 500 hours endurance test included both the European stationary and transient cycle (ESC and ETC) as well as longer stationary phases. During the test, regulated emissions (carbon monoxide, nitrogen oxides, hydrocarbons and particulate matter), the particle number distribution and the aldehydes emission were continuously measured. For comparison, tests with fossil diesel fuel were performed before and after the endurance test. During the endurance test, the engine was failure-free for 500 hours with the biogenic fuel. There were almost no differences in specific fuel consumption during the test, but the average exhaust gas temperature increased by about 15°C over the time. Emissions changed only slightly during the test.

Offering aircraft fuel efficiency improvements of 30 to 40% over the powerplants it will replace, PW2037 retrofit in the 727-200 Advanced and B-52 aircraft is attracting heightened interest. A comparison of PW2037 technical characteristics with current aircraft powerplants substantiates the improvement potential.The engine installation and modifications necessary for aircraft system compatibility do not impose significant increases in complexity or cost. The resultant improvements in aircraft capability (727 and B-52) and economic viability to airlines (7271 produce aircraft uniquely suited to today's operational requirements and constrained equipment budgets.

Flight control technology for 8000 psi has emerged almost simultaneously with new fire-resistant hydraulic fluids, such as MIL-H-83282 and CTFE. A proliferation of industry recommendations has resulted in a wide variety of mechanisms for solving associated actuator design problems, including tighter clearances, special seals, finishes, materials, and many others. As there are few common agreements on the issues, an extensive three-phase test program was undertaken to attempt to corroborate some of these approaches or suggest others that may be better or more cost effective.

An 8000 psi test program was conducted to resolve conflicts and issues surrounding the use of CTFE and MIL-H-83282 fluid with vented and unvented actuator rod seals. Each of the four possible combinations had unique problems and each responded to appropriate corrections including new backup rings designed to operate with standard clearances. It was concluded that all combinations were viable within certain limits. Advantages and disadvantages of each configuration were identified and specific recommendations made for both dynamic and static seals within the context of existing military specifications.

Shortly after World War II, as aircraft became more sophisticated and power-assist, flight-control functions became a requirement, hydraulic system operating pressures rose from the 1000 psi level to the 3000 psi level found on most aircraft today. Since then, 4000 psi systems have been developed for the U.S. Air Force XB-70 and B-1 bombers and a number of European aircraft including the tornado multirole combat aircraft and the Concorde supersonic transport. The V-22 Osprey incorporates a 5000 psi hydraulic system. The power levels of military aircraft hydraulic systems have continued to rise. This is primarily due to higher aerodynamic loading, combined with the increased hydraulic functions and operations of each new aircraft. At the same time, aircraft structures and wings have been getting smaller and thinner as mission requirements expand. Thus, internal physical space available for plumbing and components continues to decrease.

Shortly after World War II, as aircraft became more sophisticated and power-assist, flight-control functions became a requirement, hydraulic system operating pressures rose from the 1000 psi level to the 3000 psi level found on most aircraft today. Since then, 4000 psi systems have been developed for the U.S. Air Force XB-70 and B-1 bombers and a number of European aircraft including the tornado multirole combat aircraft and the Concorde supersonic transport. The V-22 Osprey incorporates a 5000 psi hydraulic system. The power levels of military aircraft hydraulic systems have continued to rise. This is primarily due to higher aerodynamic loading, combined with the increased hydraulic functions and operations of each new aircraft. At the same time, aircraft structures and wings have been getting smaller and thinner as mission requirements expand. Thus, internal physical space available for plumbing and components continues to decrease.

The 912 engine is a well known 4-cylinder horizontally opposed 4-stroke liquid-/air-cooled aircraft engine. The 912 family has a strong track record: 40 000 engines sold / 25 000 still in operation / 5 million flight hours annually. 88% of all light aircraft OEMs use Rotax engines. The 912iS is an evolution of the Rotax 912ULS carbureted engine. The “i” stands for electronic fuel injection which has been developed according to flight standards, providing a better fuel efficiency over the current 912ULS of more than 20% and in a range of 38% to 70% compared to other competitive engines in the light sport, ultra-light aircraft and the general aviation industry. BRP engineers have incorporated several technology enhancements. The fully redundant digital Engine Control Unit (ECU) offers a computer based electronic diagnostic system which makes it easier to diagnose and service the engine.

This paper describes how airplane control surface rate limiting can enable a ‘cliff-like’ onset of Pilot-Induced Oscillation, (P.I.O.) and how the danger can be erased by implementation of Conditionally Enabled Phase Compensated Rate Limiters, (PCRLs), in the design of the airplane's flight control system. The application is particularly important for large airplanes where control surface actuator sizing and the associated hydraulic system volumetric flow rate capability cannot be generously over-sized without large weight and cost penalties. It is shown that the PCRL can remain inactive during normal airplane operations where RMS control commands are relatively small thus avoiding adverse control surface response effects that have hindered earlier PCRL acceptance.

A hydrogen economy is an increasingly popular solution to lower global carbon dioxide emissions. Previous research has been focused on the economic conditions necessary for hydrogen to be cost competitive, which tends to neglect the effectiveness of greenhouse gas mitigation for the very solutions proposed. The holistic carbon footprint assessment of hydrogen production, distribution, and utilization methods, otherwise known as “well-to-wheels” carbon intensity, is critical to ensure the new hydrogen strategies proposed are effective in reducing global carbon emissions. When looking at these total carbon intensities, however, there is no single clear consensus regarding the pathway forward. When comparing the two fundamental technologies of steam methane reforming and electrolysis, there are different scenarios where either technology has a “greener” outcome.

A centrifugal pump concept was elaborated for a multiple application in future spacecrafts. Based on this concept a prototype of a small centrifugal pump was manufactured and comprehensively tested. The model pump has been approved in different test series with the fluids liquid ammonia and demineralized water. The design of the model pump was driven by the strict requirements of COLUMBUS, namely long life, noiseless operation, minimum mass and low energy consumption. Because of its modular design and as a result of selected materials of multiple compatibility, this pump is suited for the delivery of various further fluids, such as freons, hydrocarbons, propellants (hydrazine) etc.. It is also capable of pumping corrosive or toxic fluids for laboratory processes in space. The wide speed range from about 1,000 to 20,000 rpm which corresponds to a flow from about 1 to 20 l/min, permits an energy saving adaption and flow control.

As mission length and the number and complexity of payload experiments increase, so does the probability of thermodegradation contingencies (e.g. fire, chemical release and/or smoke from overheated components or burning materials), which could affect mission success. When a thermodegradation event occurs on board a spacecraft, potentially hazardous levels of toxic gases could be released into the internal atmosphere. Experiences on board the Space Shuttle have clearly demonstrated the possibility of small thermodegradation events occurring during even relatively short missions. This paper will describe the Combustion Products Analyzer (CPA), which is being developed under the direction of the Toxicology Laboratory at Johnson Space Center to provide necessary data on air quality in the Shuttle following a thermodegradation incident.

This paper attempts to assess the benefits of a unique distributed propulsion concept, known as the Multi-Gas Generator Fan (MGGF) system, over conventional turbofan engines on civilian vertical takeoff and landing (VTOL) applications. The MGGF-based system has shown the potential to address the fundamental technical challenge in designing a VTOL aircraft: the significant mismatch between the power requirements at lift-off/hover and cruise. Vehicle-level performance and sizing studies were implemented using the Grumman Design 698 tilt-nacelle V/STOL aircraft as a notional personal air vehicle (PAV), subjected to hypothetical single engine failure (SEF) emergency landing requirements and PAV mission requirements.

In recent years, a tremendous interest has spawned towards adapting Lithium-Ion battery technology for aircraft applications. Lithium-Ion technology is already being used in some military aircraft (e.g., the F-22, F-35 and the B-2) and it has also been selected as original equipment for large commercial aircraft (e.g., the Airbus A380 and Boeing B787). The advantages of Lithium-Ion technology over Lead-Acid and Nickel-Cadmium technologies are higher specific energy (Wh/kg) and energy density (Wh/L), and longer cycle life. Saving weight is especially important in aircraft applications, because it can boost fuel economy and increase mission capability. Disadvantages of Lithium-Ion technology include higher initial cost, limited calendar/float life, inferior low temperature performance, and more severe safety hazards. This paper will present a direct comparison of a 24-Volt, 28Ah Lead-Acid and a 24-volt, 28Ah Lithium-Ion aircraft battery.

Both humans and onboard radiosensitive systems (electronics, materials, payloads and experiments) are exposed to the deleterious effects of the harsh space radiations found in the space environment. The purpose of this paper is to present the space radiation environment extended to deep space based on environment models for the moon, Mars, Jupiter, and Saturn and compare these radiation environments with the earth's radiation environment, which is used as a comparative baseline. The space radiation environment consists of high-energy protons and electrons that are magnetically “trapped” in planetary bodies that have an intrinsic magnetic field; this is the case for earth, Jupiter, and Saturn (the moon and Mars do not have a magnetic field). For the earth this region is called the “Van Allen belts,” and models of both the trapped protons (AP-8 model) and electrons (AE-8 model) have been developed.

This paper presents the results of an investigation to improve design/analysis techniques for aircraft hydraulic systems. A design/analysis tool was developed by integrating control-surface commands and loads obtained from Aircraft Dynamic Simulator Software (ADSS) with an enhanced version of the HYdraulic TRansient ANalysis (HYTRAN) program. Control-surface commands and loads from an ADSS simulation of a selected maneuver were used as dynamic input to the HYTRAN program so that the hydraulic system response could be predicted throughout the maneuver. Predicted hydraulic system pressures and control-surface positions from the HYTRAN simulation of the maneuver were compared to flight-test data and were found to be in excellent agreement. The successful coalescence of the two independent software programs gives engineers a concurrent design/analysis tool that can be used to optimize hydraulic system designs during the very early stages of design.

The Shuttle orbiter currently uses hydrazine-powered APU’s for powering its hydraulic system pumps. To enhance vehicle safety and reliability, NASA is pursuing an APU upgrade where the hydrazine-powered turbine is replaced by an electric motor pump and battery power supply. This EAPU (Electric APU) upgrade presents several thermal control challenges, most notably the new requirement for moderate temperature control of high-power electronics at 132 °F (55.6 °C). This paper describes how the existing Water Spray Boiler (WSB), which currently cools the hydraulic fluid and APU lubrication oil, is being modified to provide EAPU thermal management.