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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Passive Water Assurance Delivery System

This report provides an analysis and evaluation of proposed design improvements to the internal water delivery system of NASA’s passive vegetable growing system. The VEGGIE system is designed to provide fresh vegetables and psychological benefits for astronauts aboard the International Space Station. Preliminary testing conducted by NASA for the VEGGIE system revealed flaws in the water delivery system. In these tests, the interface between the plants and the water reservoir failed to provide passive water delivery through the system both in microgravity experiments and experiments conducted on Earth. Our team’s goal was to fix this interface such that water can be passively delivered from the water reservoir to the plant rooting pillows to grow vegetative crops from seed to harvest. The system requirements outlined by NASA include: minimal total mass, on-demand passive water delivery, same dimensions as existing fixtures, minimal pressure on system to prevent leaking, zero mold growth, minimal swelling or clogging for non-flammable capillary materials, pressure stabilization between the water reservoir and plant pillow bags, even dispersal of water during initial priming, minimal bubble obstruction of capillary interface, maximum gas availability to plant roots, growth in a mixed artificial media (50:50 Arcillite:Fafard #2), and to avoid overwatering and drought conditions for plants. The major proposed design improvements are to use a single-interface capillary cord design to directly connect the water reservoir to the plant-rooting pillows, and to alter the water reservoir to model a propellant management device (PMD) in order to ensure consistent and long term watering for the VEGGIE system. The plant rooting pillow required minimal changes outside of replacing the capillary mat on the bottom of the pillow with O-ring insertion points the single-interface capillary system. - Research was conducted to determine which materials are able to uptake water through capillary action to grow Outredgeous Lettuce plants from seed to harvest. The primary requirements to be met by the team's design and evaluation of the capillary interface were: continuous passive watering for 90 days using non-flammable capillary materials and a peak water delivery rate of 30 mL/hr./0.15 m 2 . The results of the team's experimentation showed that the capillary material Nomex displayed the highest capillary water delivery potential with a maximum flow rate of 3.6 mL/hr. The Nomex capillary systems were the only capillary material to consistently grow Outredgeous Lettuce plants from seed to harvest, and displayed the highest average flow rate for multiple experiment sets. The Nomex material was previously incorporated into the VEGGIE system using a matted version of the capillary material, though the team recommends using a cord configuration in the single-interface capillary system for greater system stability. Nomex has proven to be a promising material, as it has passed both health and fire standards for use aboard the International Space Station. PMDs are made of materials that utilize surface tension and adhesive forces to overcome adverse accelerations to improve stability and ensure fluid delivery. PMDs are typically used in fuel tanks to ensure fuel delivery. The team recommends using a sponge PMD in order to mitigate bubble obstruction, decrease system weight, and ensure reliable water delivery to the capillary interface. The PMD water reservoir requires a rigid water reservoir and a vent tube for pressure stabilization. Since PMD’s cannot be tested in 1-G (Earth conditions), further theoretical modeling and testing is required for the proposed water reservoir design. The intent of this proposed system is to passively water plants in microgravity, though the technology is not limited to microgravity applications. The testing at The Ohio State University has proven that the design is highly effective on Earth, demonstrating that it could serve as a simple water delivery system in home and office applications. This would make vegetative crops more accessible in all indoor applications, thereby improving indoor air quality and occupant comfort. Additionally, this technology has great potential to be utilized in greenhouse plant production, cutting back on more sophisticated watering system energy and time requirements.

Alexandria Jensen↗
NASA NTRS2007

A Multibody Slosh Analysis for the Lunar Reconnaissance Orbiter

The Lunar Reconnaissance Orbiter (LRO) undergoes a series of thruster maneuvers to attain lunar orbit. The first of the series of lunar orbit insertion (LOI) maneuvers is crucial to the success of the mission. Therefore, it is important to characterize the disturbances acting on the spacecraft during this phase of the mission. This paper focuses on the internal disturbance force caused by fuel slosh and its impact on attitude control. During the first LOI maneuver (LOI-1), approximately 50% of the total fuel mass is used or roughly 25% of the spacecraft s wet mass, during the 38-minute burn. The forces imparted on the spacecraft from the fuel are dependent on the fill level of the two fuel tanks. During LOI-1, the fill level in both tanks varies greatly and thus so does the disturbance level caused by the fuel. It is therefore necessary to account for the time-varying mass properties of the spacecraft and the effects of the varying fuel levels during the entire 38-minute maneuver. Two simulations are developed in Mathworks s Simulink to analyze the fuel slosh effect. The first model, a baseline model, is a rigid body dynamics model where the fuel slosh is not modeled. The second is a multibody model, developed using a multibody dynamics toolbox, where each of the two fuel tanks and the remaining spacecraft body are treated as separate rigid bodies. The simulations are executed in a piece-wise fashion to account for the time-varying mass properties, and to accommodate the multibody toolbox. Disturbances caused by fuel slosh during both lunar and mission orbit insertions will be analyzed through simulation of different dynamics models. Results of the analysis will show the effects of the slosh disturbance on the spacecraft s attitude.

Shah, Neerav↗
NASA NTRS2005

Microwave and Millimeter Wave Nondestructive Evaluation of the Space Shuttle External Tank Insulating Foam

The Space Shuttle Columbia s catastrophic failure has been attributed to a piece of external fuel tank insulating SOFI (Spray On Foam Insulation) foam striking the leading edge of the left wing of the orbiter causing significant damage to some of the protecting heat tiles. The accident emphasizes the growing need to develop effective, robust and life-cycle oriented methods of nondestructive testing and evaluation (NDT&E) of complex conductor-backed insulating foam and protective acreage heat tiles used in the space shuttle fleet and in future multi-launch space vehicles. The insulating SOFI foam is constructed from closed-cell foam. In the microwave regime this foam is in the family of low permittivity and low loss dielectric materials. Near-field microwave and millimeter wave NDT methods were one of the techniques chosen for this purpose. To this end several flat and thick SOFI foam panels, two structurally complex panels similar to the external fuel tank and a "blind" panel were used in this investigation. Several anomalies such as voids and disbonds were embedded in these panels at various locations. The location and properties of the embedded anomalies in the "blind" panel were not disclosed to the investigating team prior to the investigation. Three frequency bands were used in this investigation covering a frequency range of 8-75 GHz. Moreover, the influence of signal polarization was also investigated. Overall the results of this investigation were very promising for detecting the presence of anomalies in different panels covered with relatively thick insulating SOFI foam. Different types of anomalies were detected in foam up to 9 in thick. Many of the anomalies in the more complex panels were also detected. When investigating the blind panel no false positives were detected. Anomalies in between and underneath bolt heads were not easily detected. This paper presents the results of this investigation along with a discussion of the capabilities of the method used.

Shrestha, S.↗
NASA NTRS2000

LH2 Tank Composite Coverplate Development and Flight Qualification for the X-33

In this paper, the development history for the first cryogenic pressurized fuel tank coverplates is presented along with a synopsis of the development strategy and technologies which led to success on this program. Coverplates are the large access panels used to access launch vehicle fuel tanks. These structures incorporate all of the requirements for a pressure vessel as well as the added requirement to mount all of the miscellaneous access points required for a fuel management system. The first composite coverplates to meet the requirements for flight qualification were developed on the X-33 program. The X-33 composite coverplates went from an open requirement to successful finished flight hardware with multiple unique configurations, complete with verification testing, in less than eighteen months. Besides the rapid development schedule, these components introduced several new technologies previously unseen in cryogenic composites including solutions to cryogenic shrinkage, self-supporting sealing surfaces, and highly loaded composite bosses with precision sealing interfaces. These components were proven to seal liquid hydrogen at cryogenic temperatures under maximum loading and pressure conditions.

Wright, Richard J.↗
NASA NTRS2003

Numerical Modeling, Thermomechanical Testing, and NDE Procedures for Prediction of Microcracking Induced Permeability of Cryogenic Composites

Reusable Space Vehicles will include light cryogenic composite fuel tanks that must not leak excessively even after multiple launches. Damage in cryogenic composite fuel tanks induced during manufacturing and advanced by thermomechanical cycling can accelerate leakage of the propellant. Whether the leakage exceeds tolerable levels depends on many factors, including pressure gradients, microcrack density, other damage such as delamination, connectivity of the cracks, residual stresses from manufacture, service-induced stresses from thermal and mechanical loads, and composite lay-up. Although it is critical to experimentally characterize permeability during various thermal and mechanical load histories, optimal design depends on having analytical models that can predict the effect of various parameters on performance. Our broad goal is to develop such models that are experimentally validated by destructive and non-destructive evaluation means.

Noh, Jae↗
NASA NTRS1949

Spin Investigation of a 1/29-Scale Model of the Republic XF-91 Airplane with a Conventional Tail Installed

An investigation has been conducted in the Langley 20-foot free-spinning tunnel of a 1/29-scale model of the Republic XF-91 airplane with a.conventional-tail arrangement installed. Previously, tests were made on the model with a vee tail installed. The erect spin and recovery characteristics of the model were determined for the normal loading with the wing installed at various amounts of incidence. The spin investigation also included inverted-spin tests, spin-recovery-parachute tests, tests with the center of gravity moved rearward, and tests with external fuel tanks added to the model. In addition, several tail.modifications were tested,on the model in an attempt, to improve the model's spin-recovery characteristics. The results indicate that any fully developed spin obtained on the airplane with the conventional tail installed will be satisfactorily terminated if rudder reversal is accompanied by moving the ailerons with the spin (stick right in a right spin).Decreasing the wing incidence from 6deg to -2deg should have a beneficial effect on the recovery characteristics of the airplane. Recovery characteristics by normal use of controls (full rudder reversal followed by moving the elevators down) will be satisfactory if the wing incidence,of the airplane is -2deg. Installation of external fuel tanks (with or without fuel) will have a somewhat adverse effect on the recovery characteristics of the airplane, but if the recovery technique includes movement of the ailerons to full with the spin, the spin rotation will be terminated rapidly. Varying the position of the center of gravity within the limits indicated to be possible on the airplane should not affect the recovery characteristics.

Klinar, Walter J.↗
NASA NTRS2010

Preliminary Characterization of the Altair Lunar Lander Slosh Dynamics and Some Implications for the Thrust Vector Control Design

This paper describes a conceptual design of the Thrust Vector Control (TVC) system and preliminary modeling of propellant slosh, for the Altair Lunar Lander. Altair is a vehicle element of the NASA Constellation Program aimed at returning humans to the moon. Guidance, Navigation, and Control (GN&C) is the measurement and control of spacecraft position, velocity, and attitude in support of mission objectives. One key GN&C function is the commanding of effectors that control attitude and impart delta V on the vehicle, utilizing both reaction control system (RCS) thrusters and throttling and TVC gimbaling of the vehicle main engine. Both the Altair descent and ascent modules carry fuel tanks. During thrusting maneuvers, the sloshing of liquid fuels in partially filled tanks can interact with the controlled system in such a way as to cause the overall system to be unstable. These fuel tanks must be properly placed, relative to the spacecraft's c.m., to avoid any unstable interactions. Following this will be a discussion of propellant slosh modeling work performed for the present vehicle configuration, including slosh frequency and participatory fluid mass predictions. Knowing the range of slosh mode frequencies over mission phases, the TVC bandwidth must be carefully selected so as not to excite the slosh modes at those frequencies. The likely need to increase the damping factor of slosh modes via baffles will also be discussed. To conclude, a discussion of operations procedures aimed at minimizing TVC-slosh interactions will be given.

Lee, Allan Y.↗
NASA NTRS1977

Oil cooling system for a gas turbine engine

A gas turbine engine fuel delivery and control system is provided with means to recirculate all fuel in excess fuel control requirements back to the aircraft fuel tank. This increases the fuel pump heat sink and decreases the pump temperature rise without the addition of valving other than normally employed. A fuel/oil heat exchanger and associated circuitry is provided to maintain the hot engine oil in heat exchange relationship with the cool engine fuel. Where anti-icing of the fuel filter is required, means are provided to maintain the fuel temperature entering the filter at or above a minimum level to prevent freezing thereof. In one embodiment, a divider valve is provided to take all excess fuel from either upstream or downstream of the fuel filter and route it back to the tanks, the ratio of upstream to downstream extraction being a function of fuel pump discharge pressure.

Coffinberry, G. A.↗
NASA NTRS1989

Design of an unmanned lunar cargo lander that reconfigures into a shelter for a habitation module or disassembles into parts useful to a permanent manned lunar base

NASA plans to establish a permanent manned lunar base by the first decade of the twenty-first century. It is extremely expensive to transport material from earth to the moon. Therefore, expense would be reduced if the vehicle that lands cargo on the moon could itself meet some of the material needs of establishing the lunar base. The design of a multi-functional lander that is entirely useful to the base after landing is described. Alternate designs of the overall lander configuration and possible uses of the lander and its components after landing are contained. The design solution is a lander employing the Saddlebagged Fuel Tank Configuration. After landing, its structure will be converted into a habitation module shelter that supports a protective layer of regolith. The fuel tanks will be cleaned and used as storage tanks for the lunar base. The engines and instrumentation will be saved as stock parts. Recommendations for further research and technology development to enhance future lander designs are given.

Davanay, Lisa↗
NASA NTRS2001

Safer Aircraft Possible With Nitrogen Generation

A system named On-Board Inert Gas Generation System/On-Board Oxygen Generation System (OBIGGS/OBOGS) was studied with Boeing. The study established the requirements for nitrogen purge (for fuel tank inerting and cargo compartment fire suppression) and oxygen (for passengers and crew). The nitrogen would be used for suppressing fires and fuel tank explosions on the aircraft, and the oxygen would be used for breathing gas during high-altitude or emergency operations.

Palaszewski, Bryan A.↗
NASA NTRS2011

Effects of Annealing Process on the Formability of Friction Stir Welded Al-Li Alloy 2195 Plates

Large rocket cryogenic tank domes have typically been fabricated using Al-Cu based alloys like Al-Cu alloy 2219. The use of aluminum-lithium based alloys for rocket fuel tank domes can reduce weight because aluminum-lithium alloys have lower density and higher strength than Al-Cu alloy 2219. However, Al-Li alloys have rarely been used to fabricate rocket fuel tank domes because of the inherent low formability characteristic that make them susceptible to cracking during the forming operations. The ability to form metal by stretch forming or spin forming without excessive thinning or necking depends on the strain hardening exponent "n". The stain hardening exponent is a measure of how rapidly a metal becomes stronger and harder. A high strain hardening exponent is beneficial to a material's ability to uniformly distribute the imposed strain. Marshall Space Flight Center has developed a novel annealing process that can achieve a work hardening exponent on the order of 0.27 to 0.29, which is approximately 50% higher than what is typically obtained for Al-Li alloys using the conventional method. The strain hardening exponent of the Al-Li alloy plates or blanks heat treated using the conventional method is typically on the order of 0.17 to 0.19. The effects of this novel annealing process on the formability of friction stir welded Al-Li alloy blanks are being studied at Marshall Space Flight Center. The formability ratings will be generated using the strain hardening exponent, strain rate sensitivity and forming range. The effects of forming temperature on the formability will also be studied. The objective of this work is to study the deformation behavior of the friction stir welded Al-Li alloy 2195 blank and determine the formability enhancement by the new annealing process.

Chen, Po-Shou↗
NASA NTRS1989

Shuttle-C, the Future Is Now

This video details plans for Shuttle-C, an unmanned heavy launch vehicle to carry payloads into orbit. Computer animations depict the Shuttle-C, which uses the same recoverable external boosters, external fuel tank and main orbiter engines as the existing Space Shuttles, through liftoff and entry into orbit, where it progressively jettisons the cargo shroud, external fuel tank, and nose shroud. The video also shows computer simulations of a remotely controlled orbital maneuvering vehicle positioning preassembled components of a Space Station and delivering planetary probes and lunar exploration materials to orbit.

Source record↗
NASA NTRS2016

NDARC NASA Design and Analysis of Rotorcraft - Input, Appendix 4

The NDARC code performs design and analysis tasks. The design task involves sizing the rotorcraft to satisfy specified design conditions and missions. The analysis tasks can include off-design mission performance analysis, flight performance calculation for point operating conditions, and generation of subsystem or component performance maps. The principal tasks (sizing, mission analysis, flight performance analysis) are shown in the figure as boxes with heavy borders. Heavy arrows show control of subordinate tasks. The aircraft description consists of all the information, input and derived, that denes the aircraft. The aircraft consists of a set of components, including fuselage, rotors, wings, tails, and propulsion. This information can be the result of the sizing task; can come entirely from input, for a fixed model; or can come from the sizing task in a previous case or previous job. The aircraft description information is available to all tasks and all solutions. The sizing task determines the dimensions, power, and weight of a rotorcraft that can perform a specified set of design conditions and missions. The aircraft size is characterized by parameters such as design gross weight, weight empty, rotor radius, and engine power available. The relations between dimensions, power, and weight generally require an iterative solution. From the design flight conditions and missions, the task can determine the total engine power or the rotor radius (or both power and radius can be fixed), as well as the design gross weight, maximum takeoff weight, drive system torque limit, and fuel tank capacity. For each propulsion group, the engine power or the rotor radius can be sized. Missions are defined for the sizing task, and for the mission performance analysis. A mission consists of a number of mission segments, for which time, distance, and fuel burn are evaluated. For the sizing task, certain missions are designated to be used for design gross weight calculations; for transmission sizing; and for fuel tank sizing. The mission parameters include mission takeoff gross weight and useful load. For specified takeoff fuel weight with adjustable segments, the mission time or distance is adjusted so the fuel required for the mission equals the takeoff fuel weight. The mission iteration is on fuel weight or energy. Flight conditions are specified for the sizing task, and for the flight performance analysis. For the sizing task, certain flight conditions are designated to be used for design gross weight calculations; for transmission sizing; for maximum takeoff weight calculations; and for anti-torque or auxiliary thrust rotor sizing. The flight condition parameters include gross weight and useful load. For flight conditions and mission takeoff, the gross weight can be maximized, such that the power required equals the power available. A flight state is defined for each mission segment and each flight condition. The aircraft performance can be analyzed for the specified state, or a maximum effort performance can be identified. The maximum effort is specified in terms of a quantity such as best endurance or best range, and a variable such as speed, rate of climb, or altitude.

Rotorcraft Input↗
NASA NTRS2022

NDARC - NASA Design and Analysis of Rotorcraft: Input and Data Structures - Vol 2

The NDARC code performs design and analysis tasks. The design task involves sizing the rotorcraft to satisfy specified design conditions and missions. The analysis tasks can include off-design mission performance analysis, flight performance calculation for point operating conditions, and generation of subsystem or component performance maps. Figure 1-1 illustrates the tasks. The principal tasks (sizing, mission analysis, flight performance analysis) are shown in the figure as boxes with heavy borders. Heavy arrows show control of subordinate tasks. The aircraft description (figure 1-1) consists of all the information, input and derived, that defines the aircraft. The aircraft consists of a set of components, including fuselage, rotors, wings, tails, and propulsion. This information can be the result of the sizing task; can come entirely from input, for a fixed model; or can come from the sizing task in a previous case or previous job. The aircraft description information is available to all tasks and all solutions (indicated by light arrows). The sizing task determines the dimensions, power, and weight of a rotorcraft that can perform a specified set of design conditions and missions. The aircraft size is characterized by parameters such as design gross weight, weight empty, rotor radius, and engine power available. The relations between dimensions, power, and weight generally require an iterative solution. From the design flight conditions and missions, the task can determine the total engine power or the rotor radius (or both power and radius can be fixed), as well as the design gross weight, maximum takeoff weight, drive system torque limit, and fuel tank capacity. For each propulsion group, the engine power or the rotor radius can be sized. Missions are defined for the sizing task, and for the mission performance analysis. A mission consists of a number of mission segments, for which time, distance, and fuel burn are evaluated. For the sizing task, certain missions are designated to be used for design gross weight calculations; for transmission sizing; and for fuel tank sizing. The mission parameters include mission takeoff gross weight and useful load. For specified takeoff fuel weight with adjustable segments, the mission time or distance is adjusted so the fuel required for the mission (burned plus reserve) equals the takeoff fuel weight. The mission iteration is on fuel weight or energy. Flight conditions are specified for the sizing task, and for the flight performance analysis. For the sizing task, certain flight conditions are designated to be used for design gross weight calculations; for transmission sizing; for maximum takeoff weight calculations; and for antitorque or auxiliary thrust rotor sizing. The flight condition parameters include gross weight and useful load. For flight conditions and mission takeoff, the gross weight can be maximized, such that the power required equals the power available. A flight state is defined for each mission segment and each flight condition. The aircraft performance can be analyzed for the specified state, or a maximum effort performance can be identified. The maximum effort is specified in terms of a quantity such as best endurance or best range, and a variable such as speed, rate of climb, or altitude. The aircraft must be trimmed, by solving for the controls and motion that produce equilibrium in the specified flight state. Different trim solution definitions are required for various flight states. Evaluating the rotor hub forces may require solution of the blade flap equations of motion.

NDARC↗
NASA NTRS2022

NDARC - NASA Design and Analysis of Rotorcraft: Input - Vol 3

The NDARC code performs design and analysis tasks. The design task involves sizing the rotorcraft to satisfy specified design conditions and missions. The analysis tasks can include off-design mission performance analysis, flight performance calculation for point operating conditions, and generation of subsystem or component performance maps. Figure 1-1 illustrates the tasks. The principal tasks (sizing, mission analysis, flight performance analysis) are shown in the figure as boxes with heavy borders. Heavy arrows show control of subordinate tasks. The aircraft description (figure 1-1) consists of all the information, input and derived, that defines the aircraft. The aircraft consists of a set of components, including fuselage, rotors, wings, tails, and propulsion. This information can be the result of the sizing task; can come entirely from input, for a fixed model; or can come from the sizing task in a previous case or previous job. The aircraft description information is available to all tasks and all solutions (indicated by light arrows). The sizing task determines the dimensions, power, and weight of a rotorcraft that can perform a specified set of design conditions and missions. The aircraft size is characterized by parameters such as design gross weight, weight empty, rotor radius, and engine power available. The relations between dimensions, power, and weight generally require an iterative solution. From the design flight conditions and missions, the task can determine the total engine power or the rotor radius (or both power and radius can be fixed), as well as the design gross weight, maximum takeoff weight, drive system torque limit, and fuel tank capacity. For each propulsion group, the engine power or the rotor radius can be sized. Missions are defined for the sizing task, and for the mission performance analysis. A mission consists of a number of mission segments, for which time, distance, and fuel burn are evaluated. For the sizing task, certain missions are designated to be used for design gross weight calculations; for transmission sizing; and for fuel tank sizing. The mission parameters include mission takeoff gross weight and useful load. For specified takeoff fuel weight with adjustable segments, the mission time or distance is adjusted so the fuel required for the mission (burned plus reserve) equals the takeoff fuel weight. The mission iteration is on fuel weight or energy. Flight conditions are specified for the sizing task, and for the flight performance analysis. For the sizing task, certain flight conditions are designated to be used for design gross weight calculations; for transmission sizing; for maximum takeoff weight calculations; and for antitorque or auxiliary thrust rotor sizing. The flight condition parameters include gross weight and useful load. For flight conditions and mission takeoff, the gross weight can be maximized, such that the power required equals the power available. A flight state is defined for each mission segment and each flight condition. The aircraft performance can be analyzed for the specified state, or a maximum effort performance can be identified. The maximum effort is specified in terms of a quantity such as best endurance or best range, and a variable such as speed, rate of climb, or altitude. The aircraft must be trimmed, by solving for the controls and motion that produce equilibrium in the specified flight state. Different trim solution definitions are required for various flight states. Evaluating the rotor hub forces may require solution of the blade flap equations of motion.

NDARC↗
NASA NTRS2023

NDARC - NASA Design and Analysis of Rotorcraft: Input - Appendix 3

The NDARC code performs design and analysis tasks. The design task involves sizing the rotorcraft to satisfy specified design conditions and missions. The analysis tasks can include off-design mission performance analysis, flight performance calculation for point operating conditions, and generation of subsystem or component performance maps. Figure 1-1 illustrates the tasks. The principal tasks (sizing, mission analysis, flight performance analysis) are shown in the figure as boxes with heavy borders. Heavy arrows show control of subordinate tasks. The aircraft description (figure 1-1) consists of all the information, input and derived, that defines the aircraft. The aircraft consists of a set of components, including fuselage, rotors, wings, tails, and propulsion. This information can be the result of the sizing task; can come entirely from input, for a fixed model; or can come from the sizing task in a previous case or previous job. The aircraft description information is available to all tasks and all solutions (indicated by light arrows). The sizing task determines the dimensions, power, and weight of a rotorcraft that can perform a specified set of design conditions and missions. The aircraft size is characterized by parameters such as design gross weight, weight empty, rotor radius, and engine power available. The relations between dimensions, power, and weight generally require an iterative solution. From the design flight conditions and missions, the task can determine the total engine power or the rotor radius (or both power and radius can be fixed), as well as the design gross weight, maximum takeoff weight, drive system torque limit, and fuel tank capacity. For each propulsion group, the engine power or the rotor radius can be sized. Missions are defined for the sizing task, and for the mission performance analysis. A mission consists of a number of mission segments, for which time, distance, and fuel burn are evaluated. For the sizing task, certain missions are designated to be used for design gross weight calculations; for transmission sizing; and for fuel tank sizing. The mission parameters include mission takeoff gross weight and useful load. For specified takeoff fuel weight with adjustable segments, the mission time or distance is adjusted so the fuel required for the mission (burned plus reserve) equals the takeoff fuel weight. The mission iteration is on fuel weight or energy. Flight conditions are specified for the sizing task, and for the flight performance analysis. For the sizing task, certain flight conditions are designated to be used for design gross weight calculations; for transmission sizing; for maximum takeoff weight calculations; and for antitorque or auxiliary thrust rotor sizing. The flight condition parameters include gross weight and useful load. For flight conditions and mission takeoff, the gross weight can be maximized, such that the power required equals the power available. A flight state is defined for each mission segment and each flight condition. The aircraft performance can be analyzed for the specified state, or a maximum effort performance can be identified. The maximum effort is specified in terms of a quantity such as best endurance or best range, and a variable such as speed, rate of climb, or altitude. The aircraft must be trimmed, by solving for the controls and motion that produce equilibrium in the specified flight state. Different trim solution definitions are required for various flight states. Evaluating the rotor hub forces may require solution of the blade flap equations of motion.

NDARC↗
NASA NTRS2003

Colossal Tooling Design: 3D Simulation for Ergonomic Analysis

The application of high-level 3D simulation software to the design phase of colossal mandrel tooling for composite aerospace fuel tanks was accomplished to discover and resolve safety and human engineering problems. The analyses were conducted to determine safety, ergonomic and human engineering aspects of the disassembly process of the fuel tank composite shell mandrel. Three-dimensional graphics high-level software, incorporating various ergonomic analysis algorithms, was utilized to determine if the process was within safety and health boundaries for the workers carrying out these tasks. In addition, the graphical software was extremely helpful in the identification of material handling equipment and devices for the mandrel tooling assembly/disassembly process.

Hunter, Steve L.↗