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Unified Finite-Element Method Development
Finite element method (FEM) is traditionally recognized by its versatility for
handling complicated geometries at a relative ease. This paper focuses on the development of a
FEM solver for the Compressible Fluid Dynamics problems using the Mixed Explicit Implicit (MEI) method.
The method utilizes object oriented program languages to create a unified environment for pre-processor,
the main solver, and post processor. With this approach, the user can obtain the solution to the
compressible Navier Stokes equations all in one environment, starting from the very raw data all the way
to post-processing and analysis of the results. The success of this method is illustrated in several
examples including compressible flow in laminar boundary layer shock interactions, ramjet, scramjet,
and SSME nozzles. First, the grid is automatically generated upon selection of appropriate mesh parameters.
On the same environment, the flow field parameters can also be specified before the flow solver is executed.
Upon solution convergence, the results can be displayed on the same window. The procedure can be repeated
until satisfactory results are achieved.
Aerothermal Analysis and Design Process Overview
An overview of the process and critical issues associated with aerothermal
analysis and design of supersonic and hypersonic systems will be presented. Topics will
include analysis and design tools, development of aerothermal boundary conditions using
engineering methods and the applicability of computational fluid dynamics, thermal
protection system predictive models, and airframe/internal component thermal management.
An example of the development of a hypersonic heatshield application will be provided to
demonstrate the analysis and design process.
Coupled Aeroheating/Ablation Analysis
for Re-entry Vehicles
This paper will demonstrate a new tool for analyzing an ablating material
exposed to an aeroheating environment. This tool is the result of the coupling the
thermal analysis capabilities of the Charring Material Ablation (CMA) finite difference
code with the Maneuvering Aerotherm Shape Change Code (MASCC). MASCC represents the
state-of-the-art in efficient aerothermal heating analysis. The code uses the axisymmetric
analogy and solves the integral momentum and energy boundary layer equations along
streamlines around the body. CMA was integrated into MASCC to provide a detailed 1D
in-depth thermal solution with decomposing / charring materials. The surface
temperature and ablation mass flux are explicitly coupled with the flowfield solution.
Details of the new code (ATAC3D) will be provided in the paper.
The proposed paper will demonstrate
the use of ATAC with two re-entry studies. The first
study will be a comparison with data taken for the Apollo
program. Comparison will be made with wind tunnel pressure
data, wind tunnel heat transfer data and with thermocouple
data from actual Apollo flights. The second study was
the design of the Entry, Descent, and Landing (EDL)
vehicle for the Pascal Probe. Pascal was a proposed
program to study the Martian atmosphere. ATAC was used
to provide preliminary aerodynamic coefficients and
heatshield sizing for the EDL. Comparisons were made
with CFD and DSMC calculations throughout the reentry
flight.
Advanced Thermal - Decomposition Analysis
of Rocket Motors and Other Thermal Protection Systems Using MSC.Marc-ATAS
Design of solid rocket motors requires an extensive knowledge of
the thermal behavior for reliability and optimization of the payload. Within a solid
rocket motor, a complex thermo-chemical-aerodynamic process occurs. During the launch,
the combustion of the solid propellant generates intense heat, often reaching 3600 K.
This results in a thermal decomposition of the combustion chamber housing and the nozzle
due to pyrolysis, and the ablation/erosion of the latter due to thermal, chemical,
and mechanical processes. Additionally, within the engine, radiation occurs which
is dependent upon the current geometry, and the amount of combustion that has
occurred. Recent developments have let to the solution of these problems for both
axisymmetric and three-dimensional geometries. Particular emphasis has been placed on
the efficient calculation of the thermal radiation. This paper discusses the numerical
simulation of the pyrolysis, surface energy inputs, thermal contact and radiation
calculations.
Development of a Pressure-Based CFD
Solver for All-Speed Flows on Arbitrary Polygonal Meshes in a Rule-Based Framework
Understanding the physics in combustion devices is critical in the design
and development of propulsion systems of next generation launch vehicles. CFD plays an
important role in enhancing our understanding of reacting flows in such devices. Two of
the most widely used techniques involve the solution of Reynolds-averaged Navier-Stokes
(RANS) equations using either (a) pressure-based or (b) density-based methods. Examples
of the former are the structured grid-based STREAM code and the unstructured grid-based
STREAM-UNS code. An example of a density-based code is the CHEM code; it is based on a
framework called LOCI which has been designed for intra-application coordination of
fine-grained numerical kernels and methods. Both density-based and pressure-based methods
have their advantages and disadvantages which have been well documented in the literature.
Density-based methods are typically most efficient and robust at the higher-end of the Mach
number spectrum and the pressure-based methods at the lower-end of the spectrum. Used in
conjunction, they can compliment each other to yield an optimum performance for all-speed
flows which are the norm in propulsion devices. Thus, it is desirable to develop a
computational tool that utilizes both these methods in a single framework. The framework
chosen for this is LOCI; it is an application framework designed to reduce the complexity
of assembling large-scale finite-volume applications as well as their integration with
other computational models such as heat conduction, structural deformations, etc. LOCI
utilizes a rule-based framework for the coordination of numerical value classes
constructed in C++ and includes semantics of unstructured mesh computations in rule
specifications. In the ongoing work, the LOCI system will be used to implement a
distributed memory parallelized algorithm in the context of the STREAM-UNS code. In this
paper, we will outline the salient features of the pressure-based technology embedded
in STREAM-UNS. The issues involved in transferring this technology into the LOCI
framework will be investigated. The progress made towards implementing the same in
LOCI will also be presented.
Aeroheating
Thermal Analysis Methods for Aerobraking Mars Missions
Mars missions often employ aerobraking
upon arrival at Mars as a low-mass method to gradually
reduce the orbit period from a high-altitude, highly
elliptical insertion orbit to the final science orbit.
Two recent missions that made use of aerobraking were
Mars Global Surveyor (MGS) and Mars Odyssey. Both spacecraft
had solar arrays as the main aerobraking surface area.
Aerobraking produces a high heat load on the solar arrays,
which have a large surface area exposed to the airflow
and relatively low mass. To accurately model the complex
behavior during aerobraking, the thermal analysis must
be tightly coupled to the flight mechanics, aerodynamics,
and atmospheric modeling efforts being performed during
operations. To properly represent the temperatures prior
to and during the drag pass, the model must include
the orbital solar and planetary heat fluxes. The correlation
of the thermal model to flight data allows a validation
of the modeling process, as well as information on what
processes dominate the thermal behavior. This paper
describes the thermal modeling method that was developed
for this purpose, as well as correlation for two flight
missions, and a discussion of potential improvements
to the methodology.
ZONAIR for RLV/TPS Design and Analysis
With continuing AFRL contractual support, the development of ZONA unified
hypersonic/supersonic/subsonic aerodynamic method ZONAIR and its integration into ZONA
aerothermoelastic software system including ASTROS for thermal protection system (TPS)
of RLV design/analysis was proven a successful tool. Feasibility cases studies included
a CKEM body, blunt cones and a simplified X-34 wing-body. Preceding the feasibility study,
substantial effort has been directed towards further development of a new code ZSTREAM
and using it with ZONAIR to replace the outdated streamline modules in SHVD, thus to
couple them with SHABP for aerothermoelastic applications. In the feasibility study,
the cases selected are well validated with finite-difference solutions using CFL3D.
Next, computed heat rates by applying ZONAIR with ZONA aerotheromelastic software
to X-34 through two assigned hypersonic trajectories were shown and found to agree
well with those using MINIVER. A prototypical TPS design procedure was established
using the obtained heat rates as input to MINIVER resulting in a minimum weight
TPS. With its FEM/TRIM modules ASTROS yields the trim solution and stress
distribution for a flexible X-34 at a typical trajectory point, demonstrating
the multifunctionality in the MDO capability for the present aerothermoealstic
methodology. Further, recent advances in the development of ZONAIR will also be
reported. These include: (a) Temperature mapping capability from aerodynamic to
structural grids to account for the effect of structural thermoelasticity; (b)
Development of an optimization procedure for TPS sizing using complex variable
differentiation (CVD) on MINIVER; (c) Adopt AEROHEAT methodology through
ZSTREAM metric coefficient for aerothermal/heat rate procedure and will
validate result with CFL3D/LATCH; and (d) Demonstration of automated panel
generation for aerodynamic modeling of ZONAIR using AML. In the forthcoming
paper, we will present detailed solutions through example cases selected
for these recent developments.
Windmill Powered Cooling for Avionics
As part of an international project, in the early '90s FIMAC S.p.A.
developed a ram air powered cooling system (Cooler) able to keep below 70 °C the cold
walls of avionic equipment on board of a fighter plane Pod, without external power input
and providing refrigerating power up to 1500 W. The Cooler obtains energy from the airflow
gathered at the Pod ram intake that crosses the air path of the Cooler. In its path, the
air drives a Ram Air Turbine (RAT) that is the windmill prime mover for the refrigerant
circuit, where the hybrid cooling capabilities of expanding air and a reverse Rankine
cycle, based on organic fluid refrigerant, are exploited to a maximum. A Multivortexâ
Compressor driven by the RAT provides the gas compression phase of the refrigerant cycle.
A compact heat exchanger Condenser rejects into the RAT exhaust air the heat from the
refrigerant gas condensation phase. Useful heat subtraction (liquid refrigerant expansion
and evaporation phases) takes place in a set of Cold Walls located in the frame of the
payload avionic boxes. Peculiar features of the Cooler, beyond conventional refrigerators,
are:
- Lower condensation temperature in the correspondent speed condition. Because of the mechanical power supplied to the RAT, the incoming air cools itself and performs a thermal exchange at a temperature lower than the stagnation temperature TRAM (this is the so-called "hybrid" cycle effect).
- Liquid Booster Pump (on the same shaft of the Compressor) delivering sub-cooled liquid flow to the Cold Walls.
- Trimming circuit of the refrigerant circulating volume (circuit charge) composed by a Tank with an electrically powered Pump and solenoid valve. Flooding or draining the Condenser, transferring fluid to and from the tank, determines the charge control.
- Air outlet nozzle control device, composed by a louver valve and hydromechanical Actuator, operated by the same refrigerant. This device modulates the air outlet area to keep the airflow and the RAT speed controlled and adequate to the operating conditions.
The operation of the Cooler involves
the combination of three functionally interacting loops:
- Air circuit for power extraction and heat rejection.
- Vapour cycle based on organic refrigerant for two-phase cooling of the Cold Walls.
- Control circuit for the determination of the active charge of the refrigerant.
The above operating loops, that enable the Cooler to operate
over the aircraft flight envelope, are intrinsically
stable and widely self-adjusting, while interaction
effects are minimised to avoid hunting. The main control
effect, the actuation of the air circuit control louver
valve, is self-regulating by action of a refrigerant
fluid operated, pressure-controlled actuator. An integrated
Electronic Control Unit receives pressure signals from
the refrigerant cycle and operates the refrigerant charge
control to optimize the working condition, delivering
in the same time status information to the aircraft.
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