A STUDY ON OPTIMAL TRAJECTORY PROGRAMMING
Optimal trajectory geometry and thrust programming
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Large Engine Technology Program. Task 21: Rich Burn Liner for Near Term Experimental Evaluations
The objective of the task reported herein, which was conducted as part of the NASA sponsored Large Engine Technology program, was to define and evaluate a near-term rich-zone liner construction based on currently available materials and fabrication processes for a Rich-Quench-Lean combustor. This liner must be capable of operation at the temperatures and pressures of simulated HSCT flight conditions but only needs sufficient durability for limited duration testing in combustor rigs and demonstrator engines in the near future. This must be achieved at realistic cooling airflow rates since the approach must not compromise the emissions, performance, and operability of the test combustors, relative to the product engine goals. The effort was initiated with an analytical screening of three different liner construction concepts. These included a full cylinder metallic liner and one with multiple segments of monolithic ceramic, both of which incorporated convective cooling on the external surface using combustor airflow that bypassed the rich zone. The third approach was a metallic platelet construction with internal convective cooling. These three metal liner/jacket combinations were tested in a modified version of an existing Rich-Quench-Lean combustor rig to obtain data for heat transfer model refinement and durability verification.
Hautman, D. J.↗
Spacelab Phase B study environmental control system component handbook
Descriptions are given of candidate components for the Phase B Study of the Spacelab environmental control system. Most of the hardware was defined for the baseline ECS design concept. A low cost design approach was followed, with most of the components being selected from the Apollo and Skylab Programs.
Burns, R. A.↗
A Monte Carlo error analysis program for near-Mars, finite-burn, orbital transfer maneuvers
A computer program was developed which performs an error analysis of a minimum-fuel, finite-thrust, transfer maneuver between two Keplerian orbits in the vicinity of Mars. The method of analysis is the Monte Carlo approach where each off-nominal initial orbit is targeted to the desired final orbit. The errors in the initial orbit are described by two covariance matrices of state deviations and tracking errors. The function of the program is to relate these errors to the resulting errors in the final orbit. The equations of motion for the transfer trajectory are those of a spacecraft maneuvering with constant thrust and mass-flow rate in the neighborhood of a single body. The thrust vector is allowed to rotate in a plane with a constant pitch rate. The transfer trajectory is characterized by six control parameters and the final orbit is defined, or partially defined, by the desired target parameters. The program is applicable to the deboost maneuver (hyperbola to ellipse), orbital trim maneuver (ellipse to ellipse), fly-by maneuver (hyperbola to hyperbola), escape maneuvers (ellipse to hyperbola), and deorbit maneuver.
Green, R. N.↗
Experimental and Modeling Study of the Burning of an Ethanol Droplet in Microgravity
The microgravity ethanol droplet combustion experiments were performed aboard the STS-94/MSL-1 Shuttle mission within the Fiber-Supported Droplet Combustion-2 (FSDC-2) program. The burning histories and flame standoffs for pure ethanol and ethanol/water droplets were obtained from the images recorded with two 8 mm videocameras. The obtained results show that average gasification rate is related to the initial droplet size in a manner similar to n-alkanes and methanol and consistent with the results of Hara and Kumagai and the data taken recently in the NASA-Lewis 2.2 s droptower. A transient, moving finite-element chemically reacting flow model applied previously to sphero-symmetric combustion of methanol, methanol/water, n-alkane, and n-alkane binary mixture droplets was adopted for the problem of ethanol droplet combustion. The model includes detailed description of gas-phase reaction chemistry and transport, a simplified description of liquid phase transport, and non-luminous radiative heat transfer. Gas-phase chemistry was described with the detailed reaction mechanism of Norton and Dryer, which consists of 142 reversible elementary reactions of 33 species. Another recently published reaction mechanism of high-temperature ethanol oxidation was also considered. The model predictions were found to compare favorably with the experimental data. The model analysis also indicates that water condensation in the case of ethanol has smaller effect on average droplet gasification rate as compared with previously studied methanol cases. This effect is explained by non-ideal (azeotropic) behavior of binary ethanol-water mixtures. Further analysis of computational results and ethanol droplet radiative extinction behavior will be discussed.
Kazakov, Andrei↗
Simulation of mixing in the quick quench region of a rich burn-quick quench mix-lean burn combustor
A computer program was developed to study the mixing process in the quick quench region of a rich burn-quick quench mix-lean burn combustor. The computer program developed was based on the density-weighted, ensemble-averaged conservation equations of mass, momentum (full compressible Navier-Stokes), total energy, and species, closed by a k-epsilon turbulence model with wall functions. The combustion process was modeled by a two-step global reaction mechanism, and NO(x) formation was modeled by the Zeldovich mechanism. The formulation employed in the computer program and the essence of the numerical method of solution are described. Some results obtained for nonreacting and reacting flows with different main-flow to dilution-jet momentum flux ratios are also presented.
Shih, Tom I.-P.↗
Status of the CNRS-LCSR program on high pressure droplet vaporization and burning
Depending on the surrounding flow and thermodynamic conditions, a single droplet may experience several gasification regimes, ranging from the envelope flame regime to pure vaporization. In practical situations, such as rocket propulsion or diesel combustion, the size distribution of droplets is, at best, bimodal, so that the possibility exists for the simultaneous presence of various regimes. For example, very small droplets are transported by the gas phase with zero relative velocity. This picture validates then the spherical symmetry hypothesis applied to the droplet and to the diffusion flame enveloping it. On the other hand, for larger droplets, a relative velocity exists due to drag forces. The most important influence of forced convection on droplet burning is the possibility to extinguish globally the envelope flame, or to establish a flame stabilized in the wake region. The burning rates of these regimes differ strongly. The characteristic time of droplet gasification is also influenced by the surrounding pressure and temperature. A parametric investigation of single droplet burning regimes is then helpful in providing the necessary physical ideas for sub-grid models used in spray combustion numerical prediction codes. The CNRS-LCSR experimental program on droplet vaporization and burning deals with these various regimes: stagnant and convective monocomponent droplet burning convective mono and bicomponent droplet vaporization; high temperature convective mono and biocomponent droplet vaporization; burning regimes of hydrazine and hydroxyl-ammonium-nitrate based monopropellant droplets and the vaporization regimes of liquid oxygen droplets. Studies on interacting droplets and on liquid aluminum droplets will start in the near future. The influence of high pressure is a common feature of all these studies. This paper summarizes the status of the CNRS-LCSR program on the effects of high pressure on monocomponent single droplet burning and vaporization, and some recent results obtained under normal and reduced gravity conditions with suspended droplets are presented. In the work described here, parabolic flights of an aircraft is used to create a reduced gravity environment of the order of 10(exp -2) g.
Chauveau, Christian↗
Design and performance analysis of solid-propellant rocket motors using a simplified computer program
An analysis and a computer program are presented which represent a compromise between the more sophisticated programs using precise burning geometric relations and the textbook type of solutions. The program requires approximately 900 computer cards including a set of 20 input data cards required for a typical problem. The computer operating time for a single configuration is approximately 1 minute and 30 seconds on the IBM 360 computer. About l minute and l5 seconds of the time is compilation time so that additional configurations input at the same time require approximately 15 seconds each. The program uses approximately 11,000 words on the IBM 360. The program is written in FORTRAN 4 and is readily adaptable for use on a number of different computers: IBM 7044, IBM 7094, and Univac 1108.
Sforzini, R. H.↗
The F/A-18 external burning flight test
A flight test program was undertaken to demonstrate the feasibility of obtaining external burning (EB) data in a flight environment, and to address the question of facility interference in ground test external burning data. Results showed that external burning was effective at reducing transonic base drag in a flight environment having dynamically changing conditions. The flight test program demonstrated that external burning is not just a laboratory phenomenon but is a viable technology for aerospace vehicles.
Yungkurth, C.↗
Low Stretch PMMA Burning in Microgravity: Status of the Ground-Based Program and New ISS Glovebox Experiment SALSA
The objective of this ground-based program is to study low stretch diffusion flames burning PMMA as the solid fuel to determine the relationship between buoyant low stretch burning in normal gravity and forced flow low stretch burning in microgravity. The low stretch is generated in normal gravity by using the buoyant convection induced by burning the bottom of a large radius of curvature sample. Low stretch is also generated using the Combustion Tunnel drop tower rig (2.2 and 5.2 second facilities), which provides a forced convective low velocity flow past smaller radius of curvature samples. Lastly, an ISS glovebox investigation is being developed to study low stretch burning of PMMA spheres to obtain long duration testing needed to accurately assess the flammability and burning characteristics of the material in microgravity. A comparison of microgravity experiment results with normal gravity test results allows us to establish a direct link between a material's burning characteristics in normal gravity (easily measured) with its burning characteristics in extraterrestrial environments, including microgravity forced convective environments. Theoretical predictions and recent experimental results indicate that it should be possible to understand a material's burning characteristics in the low stretch environment of spacecraft (non-buoyant air movement induced by fans and crew disturbances) by understanding its burning characteristics in an equivalent Earth-based low stretch environment (induced by normal gravity buoyancy). Similarly, Earth-based stretch environments can be made equivalent to those in Lunar- and Martian-surface stretch environments (which would induce partial-gravity buoyancy).
Olson, S. L.↗
Hybrid composites that retain graphite fibers on burning
A laboratory scale program was conducted to determine fiber release tendencies of graphite reinforced/resinous matrix composites currently used or projected for use in civil aircraft. In the event of an aircraft crash and burn situation, there is concern that graphite fibers will be released from the composites once the resin matrix is thermally decomposed. Hybridizing concepts aimed at preventing fiber release on burning were postulated and their effectiveness evaluated under fire, impact, and air flow during an aircraft crash.
House, E. E.↗
A study of the role of pyrotechnic systems on the space shuttle program
Pyrotechnic systems, high burn rate propellant and explosive-actuated mechanisms, have been used extensively in aerospace vehicles to perform a variety of work functions, including crew escape, staging, deployment and destruction. Pyrotechnic system principles are described in this report along with their applications on typical military fighter aircraft, Mercury, Gemini, Apollo, and a representative unmanned spacecraft. To consider the possible pyrotechnic applications on the space shuttle the mechanical functions on a large commercial aircraft, similar in scale to the shuttle orbiter, were reviewed. Many potential applications exist for pyrotechnic system on the space shuttle, both in conventional short-duration functions and in longer duration and/or repetitive type gas generators.
Lake, E. R.↗
A comparison of burn-in and bake as semiconductor screening techniques for the nimbus spacecraft program
Comparison of high temperature storage bake and operating burn-in as screening techniques for semiconductor devices used in Nimbus satellite
NIMBUS SATELLITE↗
Space tug geosynchronous mission simulation
Near-optimal three dimensional trajectories from a low earth park orbit inclined at 28.5 deg to a synchronous-equatorial mission orbit were developed for both the storable (thrust = 28,912 N (6,500 lbs), I sub sp = 339 sec) and cryogenic (thrust = 44,480 N (10,000 lbs), I sub sp = 470 sec) space tug using the iterative cost function minimization technique contained within the modularized vehicle simulation (MVS) program. The finite burn times, due to low thrust-to-weight ratios, and the associated gravity losses are accounted for in the trajectory simulation and optimization. The use of an ascent phasing orbit to achieve burnout in synchronous orbit at any longitude is investigated. The ascent phasing orbit is found to offer the additional advantage of significantly reducing the overall delta velocity by splitting the low altitude burn into two parts and thereby reducing gravity losses.
Lang, T. J.↗
Human Mars Mission Performance Crew Taxi Profile
This timeline was generated on the Integrated Mission Program (IMP). All burn events over 2 seconds are finite with IMP solving a two point boundary value setup for begin burn time, burn time and control angles. Perigee and apogee shown above are mean orbital values. Significant events are listed. Each finite thrust event has two lines. The first is the beginning time showing the initial conditions, thrust and ISP used. The second has the end burn conditions and the delta v and time of burn. This case is an abort from the 750 x 750 phasing abort, using the taxi's main engines. An abort using the Reaction Control System (RCS) was also investigated but required a large increase in RCS propellant and was abandoned.
Duaro, Vince A.↗
Aircraft Engine Technology for Green Aviation to Reduce Fuel Burn
The NASA Fundamental Aeronautics Program Subsonic Fixed Wing Project and Integrated Systems Research Program Environmentally Responsible Aviation Project in the Aeronautics Research Mission Directorate are conducting research on advanced aircraft technology to address the environmental goals of reducing fuel burn, noise and NOx emissions for aircraft in 2020 and beyond. Both Projects, in collaborative partnerships with U.S. Industry, Academia, and other Government Agencies, have made significant progress toward reaching the N+2 (2020) and N+3 (beyond 2025) installed fuel burn goals by fundamental aircraft engine technology development, subscale component experimental investigations, full scale integrated systems validation testing, and development validation of state of the art computation design and analysis codes. Specific areas of propulsion technology research are discussed and progress to date.
Hughes, Christopher E.↗
Development of advanced diagnostics for characterization of burning droplets in microgravity
Diagnostic techniques currently used for microgravity research are generally not as advanced as those used in earth based gravity experiments. Diagnostic techniques for measuring the instantaneous radial temperature profile (or temperature gradients) within the burning droplet do not exist. Over the past few years, Aerometrics has been researching and developing a rainbow thermometric technique for measuring the droplet temperatures of burning droplets. This technique has recently been integrated with the phase Doppler interferometric technique to yield a diagnostic instrument that can be used to simultaneously measure the size, velocity, and temperature of burning droplets in complex spray flames. Also, the rainbow thermometric technique has been recently integrated with a point-diffraction interferometric technique for measuring the instantaneous gas phase temperature field surrounding a burning droplet. These research programs, apart from being very successful, have also helped us identify other innovative techniques for the characterization of burning droplets. For example, new techniques have been identified for measuring the instantaneous regression rate of burning droplets. Also, there is the possibility of extracting the instantaneous radial temperature distribution or the temperature gradients within a droplet during transient heating. What is important is that these diagnostic techniques have the potential for making use of inexpensive, light-weight, and rugged devices such as diode lasers and linear CCD arrays. As a result, they can be easily packaged for incorporation into microgravity drop-test and flight-test facilities. Furthermore, with the use of linear CCD arrays, data rates as high as 10-100 kHz can be easily achieved. This data rate is orders of magnitude higher than what is currently achievable. In this research and development program, a compact and rugged diagnostic system will be developed that can be used to measure instantaneous fuel droplet diameter, droplet regression rate, and the droplet internal temperature profiles or gradients at very high data rates in microgravity experiments.
Sankar, Subramanian↗
OPTIM: Computer program to generate a vertical profile which minimizes aircraft fuel burn or direct operating cost. User's guide
A profile of altitude, airspeed, and flight path angle as a function of range between a given set of origin and destination points for particular models of transport aircraft provided by NASA is generated. Inputs to the program include the vertical wind profile, the aircraft takeoff weight, the costs of time and fuel, certain constraint parameters and control flags. The profile can be near optimum in the sense of minimizing: (1) fuel, (2) time, or (3) a combination of fuel and time (direct operating cost (DOC)). The user can also, as an option, specify the length of time the flight is to span. The theory behind the technical details of this program is also presented.
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