AIAA-92-3721 Evaluation of Thermal Management for a Mach 5.5 Hypersonic Vehicle

James A. Gasner and Ronald C. Foster General Dynamics Corporation Fort Worth, Texas 76101

Clay Fujimura Wright Laboratory Wright-Patterson AFB, Ohio 45433

AI AAISA €/AS M €/AS € E 28th Joint Propulsion

Conference and Exhibit July 6-8, 1992 / Nashville, TN

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For permission to copy or republish, contact the American institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

AIM-92-3721

EVALUATION OF THERMAL MANAGEMENT FOR A MACH 5 . 5 HYPERSONIC VEHICLE

4

James A. Gasner Thermodynamic Analysis Engineer-Senior

Fort Worth, Texas 76101

Ronald C. Foster Thermodynamics Analysis Engineering Specialist, Sr.

Fort Worth, Texas 76101

and

Clay Fuj imura aerospace Engineer

Wright Laboratory (WL/FIMM) Wright-Patterson AFB, Ohio 4 5 4 3 3

General Dynamics Corporation, Fort Worth Division

General Dynamics Corporation, Fort Worth Division

Abstract

This paper discusses an analytical evaluation of the thermal management (TM) system for a Mach 5 . 5 hypersonic air- breathing vehicle. Endothermic fuel is utilized as a heat sink as well as the propulsion source for the aircraft. The heat-absorbing reaction that the fuel undergoes in the presence of a catalyst permits the TM system to be integrated within the fuel supply system. Airframe and engine cooling can be achieved by utilizing the fuel supply system as the available heat sink to cool high temperature air for use in film cooling engine and nozzle hot

4 structures. Transferring these heat loads to the fuel as it flows to the engine allows the heat to be rejected at the engine combustor where the fuel is burned: thereby, dumping the heat overboard into the exhaust stream. The thermal modelling techniques employed to analyze the integral fuel supply/thermal management system are explained. Typical thermal study results are presented in terms of fuel system temperature profiles throughout a Mach 5 . 5 design mission. Study results indicate the feasibility of developing TM systems for hypersonic vehicles.

Introduction

Successful development of a Mach 5 . 5 hypersonic vehicle requires the evaluation and integration of a thermal management (TM) system as an essential part of the airframe and propulsion system design. At hypersonic speeds aerodynamic heating of the aircraft and engine/nozzle cooling requirements are significant heat loads for the TM system.

Coovriaht O 1992 bv General Dvnamics

Utilization of the fuel as a heat sink becomes necessary to meet vehicle cooling requirements. Therefore, the thermal management system needs must be recognized in the overall fuel management optimization to balance engine fuel flow demand, engine and airframe heat loads, and available fuel heat sink. This must be accomplished while maintaining the fuel temperatures below limits imposed by the fuel chemical stability and/or the fuel supply system thermal requirements.

Under a contracted effort funded by Wright Laboratory from 1988 to 1990, General Dynamics-Fort Worth Division conducted a study to identify technology requirements and define a conceptual design of an integrated airframe/propulsion system for a Mach 5 . 5 hypersonic vehicle’. As part of this program, a vehicle thermal management concept was also defined and. evaluated. Thermal analysis was performed on an open- loop high temperature air system in which the fuel cools engine fan or ram air so it can be utilized for film cooling of engine and nozzle hot structures. Similar TM studies have been conducted by NASA-Langley and have determined the feasibility of operational hypersonic vehicles using non- cryogenic fuel and a secondary closed loop cooling system2.

.I - Corporation. hlb1is;ed by the American Institute of Aeronautics and Astronautics, Inc. with permission.

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Aero-Heatinq Survey of Tvpical Vehicle

The vehicle design mission (Mach and altitude profile versus time) is presented in Figure 1. As an initial step in developing a TM system, an aero-heating survev of the vehicle is conducted to evaluate active coolina reauirements for the vehicle structure 6y calculating the structural temperature extremes. These extreme temperatures can be predicted by determining the radiation equilibrium temperatures at various aircraft locations. The General Dynamics Aerodynamic/Structural Heating with ABlation (ASHAB) computer code was utilized to predict radiation equilibrium temperatures over the vehicle during Mach 5.5 cruise conditions3. The general shape of the vehicle is represented by combinations of the simple geometric shapes that the program considers. Geometries such as spheres are used to model the vehicle nose: swept cylinders, to model wing leading edges; flat plates, wedges, or cones, to simulate various fuselage locations. The radiation eauilibrium temperatures over the vehicle are-predicted from flowfield correlations for these simplified shapes. Laminar and turbulent aerbdynamic heating is calculated using Eckert reference enthalpy techniques.

. . . . . . . . . . 6 . 0 1 : OYIDO"*D C"YI5T ' : . . . . : : ~ : I

2

Figure 1. Hypersonic Vehicle Design Mission contains Mach 5 . 5 cruise Leg.

The predicted radiation equilibrium temperatures over the aircraft during the outbound cruise (Mach 5.5, 85K ft.) are shown in Figure 2 . The vehicle is assumed to fly at an angle of attack of 7' in Standard Day atmospheric conditions and has a surface emissivity of 0.8 for radiation purposes. ASHAB results predict fuselage temperatures ranging from 750'F to 120OoF, tail and wing leading edge temperatures from 1450'F to 1500°F. and vehicle nose temperatures reaching approximately 1800'F.

Figure 2. Vehicle Structural Temperatures Predicted from the ASHAB L' Aeroheating Code.

The airframe structural active cooling requirements are assumed to be minimal for the present thermal management studies since titanium-aluminide composites, which are proposed for use on these vehicles based on their availability, exhibit the required strength at elevated temperatures up to 1800'F. Shock wave interaction effects generate extreme localized heating along the engine inlet cowl lip. To decrease the thermal stress levels, the cowl lip is actively cooled to reduce the temperature gradients in this region. Analysis of the cowl lip cooling requirement is presented in a later section. The evaluation of the active cooling for the entire inlet structure was beyond the scope of this analysis but presents many challenges for development of moveable structures that can handle temperatures up to 2500'F. Carbon- carbon composites were assumed to be used over the present vehicle inlet with no active cooling required. v

Although the structural materials of the vehicle are assumed to be able to withstand these extreme temperatures, the heating effect on the bulk fuel temperature needs to be evaluated. As the on-board fuel temperature increases, the available heat sink capacity of the fuel is diminished. It

4 is important to keep the fuel tanks insulated from the external aircraft structures so the fuel may be utilized as an effective TM heat sink.

s u B W m *

Thermal Manasement Model Development

Thermal management techniques for hypersonic vehicles have been studied extensively by General Dynamics under Independent Research and Development (IRAD) programs. During these IRAD programs, a Hypersonic Vehicle Thermal Management (HVTM) methodology was developed. The results of this work and the methodoloqies developed

E3 of reaction products) occurs at approximately 1000°F. Temperature

increased cooling capacities are essential to successful thermal management of hypersonic vehicles. Methylcyclohexane (MCH) was chosen to fuel the hypersonic vehicle developed in this program. MCH is an endothermic hydrocarbon fuel which has a heat sink capacity approximately ten times that of the more traditional JP-type hydrocarbon fuels. Cryogenic Methane and Hydrogen also have increased heat sink capacities, but the increased logistic requirements associated with handling/ storage and problems related to fuel boil off of cryogenic fuels eliminate these as practical fuels for this application.

The increased heat sink capacity of MCH results from the endothermic chemical reaction initiated by a Platinum-based catalyst. The reaction products formed from this reaction are toluene and hydrogen as shown below:

-mm ,,EA, /PA, catalytic reactors are needed in the TM U"l*N(irn ?m'P system and their size requirements6. _Cl/ma

<*-Me IAII1rnM"

Fuel Heat Sink Characterization- Methvlcvclohexane(MCH~

The Air Force has been developing endothermic fuels as @ea way of orovidina Figure 4. Temperature-Enthalpy Properties extremely high heat sink onlboard- aircraff and missile^"^. Since fuel is to be

L/ utilized as the heat sink, fuels with

of Methylcyclohexane Display the Favorable Heat Absorbing Quality of an Endothermic Fuel.

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Aero-Heatins Effects on Tank Bulk Fuel TemDeratures

AS previously stated, it is necessary to insulate the fuel tanks to reduce the impact of aerodynamic heating on bulk fuel temperatures. Figure 5 illustrates the various heat transfer mechanisms which affect the bulk fuel temperature. The thermal model created accounts for insulation of the fuel tanks with a high temDerature fsteadv-state temDerature limit of -18000~) , 'flexihe blanket-type, Min-KB insulation produced by the Manville Aerospace company'. The fuel tank model incorporates the external aero-heating along the top and bottom of the vehicle as determined from the ASHAB computer code for the various fuel tank locations. External radiation cooling to the associated "clear- sky" temperatures, the conduction through the insulation blanket and fuel tank wall, and the internal convection between the tank wall and the fuel are also simulated. This internal convection is a function of the wetted surface area of the tank wall and varies as fuel flows out or into the tank. The heat transfer associated with fuel flowing between fuel tanks is also modelled. Fuel transfer is established from the fuel management schedule which empties the tanks in a sequence which controls the center of gravity of the vehicle. Radiation heat transfer between the upper and lower tank wall surfaces is also incorporated in the model in a simplified manner assuming the fuel acts as a non-participating radiation medium.

Figure 5. Insulation of Fuel Tanks Reduces Aeroheating Effects on Tank Fuel Temperatures.

The thermal management strategy for the hypersonic vehicle utilizes the fuel supply system, i.e. fuel flowing to the engine, as a heat sink to meet the coolins resuirements

these different cooling requirements vary throughout the mission duration and incorporate these time varying heat loads into the SINDA thermal model. This allows prediction of the fuel system temperature profiles and produces a "picture" of the

Aircraft Subsvstems Coolinq Reouirements

thermal management status of the vehicle. L'

The aircraft subsystems cooling requirements are based on I R A D work previously conducted by General Dynamics in development of a vapor-cycle-based ECS system for a Mach 5.5 hypersonic interceptor vehicle. At Mach 5.5, a 40 kW cabin cooling requirement is estimated as a result of aerodynamic heating of the vehicle crew station. The following aircraft electrical requirements are utilized in the analysis of subsystem cooling requirements:

KLQ AVG ELECTRICAL REO, kW

Flight Controls 110 Avionics Systems 20 Fuel Management System 35 Advanced Radar 60

Total Avionics 225 kW (165 kW-with Radar off)

It is assumed 100% of the total avionics electrical load is converted to heat that must be removed by the ECS system and an 8 0 % efficient electrical generator supplies this power. Therefore, the avionics plus generator cooling requirement is 281 kW when

Figure 6 illustrates the ECS concept considered for subsystem cooling which employs cascaded vapor compression- refrigeration cycles dissipating heat to the fuel as a heat sink. Compressor efficiencies are assumed to be 70% with compressor power also supplied from the electrical generator.

the radar is operating. L'

of the airframe and engine whiie maintaining the fuel temperatures below stability Figure 6 . Vehicle Subsystem Cooling Concept -, limits. It is necessary to predict how Implements CascadedVaporCycies.

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Analysis results for the proposed ECS system are presented in Figure 7. The ECS heat rejection in kW is plotted as a function of fuel flow rate assuming 150'F inlet fuel temperature. Two curves are shown for avionics loads with and without - the 60 kW radar operating. Note how Figure 7 indicates that the heat rejection to the fuel increases at lower fuel flow rates. This effect is caused by the resulting increase in the fuel temperature in the ECS heat exchangers at lower fuel flow rates, which results in higher pressure ratios (and hence higher power requirements) for the ECS compressors. These subsystem heat load characteristics as a function of fuel flow rate are input into the thermal management model to account for aircraft subsystem heating of the fuel.

W

Figure 7. Vehicle subsystem Heat Rejection Varies as a Function of Fuel Flow Through the Heat Exchanger.

Cowl Lip Coolinu Reuuirements

Although the aero-heating survey indicated no aircraft structures required active cooling, the shock structure of the engine inlet at Mach 5.5 indicates the need for active cooling of the cowl lip of the engine inlet. During the outbound cruise of the mission, the external shock waves generated from the variable external inlet ramps (inlet compression surfaces) are focused near the cowl lip to reduce drag. These shock interactions cause supersonic "scrubbing" on the cowl lip and intensify

the localized aerodynamic heating effects on the structure. The thermal management model, therefore, includes a second airframe heat exchanger to transfer the cowl lip heat load to the fuel. The approach used to predict the cowl lip cooling requirement is explained in the following paragraphs.

The effects of shock impingement heating on the engine inlet cowl lip were evaluated based on experimental results of shock wave interference heating on a cylindrical leading edge*. Peak heat transfer rates of 5 to 10 times the undisturbed values for stagnation point heating were observed dependent on the strength of the impinging shock. For tests conducted at Mach 6 . 5 , the referenced study presents the variation in the ratio of shock interference heating to undisturbed stagnation point heating (q/q,) along the circumference of a cylinder exposed to a Type I V shock interaction (an oblique shock intersects the nearly normal part of the bow shock from a blunt leading edge). Based on the test results presented, a simplified equation was formulated as:

Uq, = 5 cos e Where theta (e) equals the angle measured around the circumference of the cylindrical leading edge (0° corresponds to the stagnation point of the leading edge).

Integrating the above equation from theta equals -90' to +90° and multiplying by the length of the cowl lip results in the following equation which was used to approximate the shock interaction heating on the cowl lip:

Q c w t t ip = 10 qo r L where: %= undisturbed stagnation point

heating rate

edge r = radius of cylindrical leading

L = length of cylindrical cowl lip

The undisturbed stagnation heating, qo, is calculated using the ASHAB aero-heating cod% for a swept cylinder with a 0' sweep angle and assuming cowl lip structural temperatures are limited to 1500'F. The cowl lip geometry is approximated with a cylinder having a 0.5" radius. During the outbound cruise (Mach 5.5, 85000 ft.), the total cowl lip cooling for four engine inlets depends on the overall cowl lip length and is estimated as 7150 Btu/min for four-50 inch width inlets. Cowl lip cooling is presumed necessary only when stagnation temperatures are greater than 1500'F and the thermal management model accounts for the variation in this heat load as a function of mission time.

Enaine and Nozzle Heat Loads

A s displayed in the Figure 3 schematic, fuel leaving the cowl lip heat exchanger passes through an engine fuel pump where it is pumped up to pressures ranging from 700

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to 1000 psia. These pressures are above the critical pressures for MCH (504 psia) and prevent vaporization of MCH as it flows through two engine system heat exchangers. The first of these heat exchangers cools engine components and the second cools engine and nozzle hot structures and requires integration of a catalyst to take advantage of the endothermic qualities of MCH fuel. Incorporation of engine and nozzle heat loads into the thermal management model resulted from an effort between General Dynamics and GE Aircraft Engines (GEAE) . Engine and nozzle cooling concepts proposed by GEAE utilize high pressure air (either engine compressor bleed air from the turbofan or ram air extracted from the engine inlet during ramjet operation) that is cooled by fuel-air heat exchangers and is then compressed and flows over hot components and structures to meet engine and nozzle cooling requirement^^.'^. This study assumes 1.25% of the engine inlet airflow must be cooled to 700'F to meet the cooling requirements of the engine components (actuators and engine oil reservoir). The GEAE approach for predicting engine and nozzle hot structures cooling requirements during ramjet mode entails determining cooling air flow rates required to: (a) cool the turbo-machinery, (b) cool the nozzle walls, and (c) supply cooled high pressure air to two ram air turbines coupled to a cooling air pump and a secondary power generator. The air is assumed to be supplied from the engine inlet at ram temperatures and requires cooling by the reactor heat exchanger to 1150'F.

ComDarison of Vehicle Heat Loads Durinq Miss ion

The various vehicle heat loads, as calculated by the techniques discussed in the preceding paragraphs, are compared in Figure 8 throughout the entire mission. The

Figure 8 . Comparison of Vehicle Heat Loads Throughout Mission Indicates the Dominance of the Engine/Nozzle Hot Structures Cooling Require- ments.

plotted values represent the total heat transferred to the fuel at each of the four heat exchangers shown in Figure 3 . During the outbound cruise of the mission (approximately 30-40 minutes), the engine/nozzle hot structures heat load is the most dominant and is approximately 15 times that of the engine components, 85 times that of the aircraft subsystems, and 350 times that of the cowl lip cooling requirements. During the return cruise at Mach 3 , this engine/nozzle structural cooling requirement is eliminated since exhaust gas temperatures drop below structural limitations (2700 OF) and the engine is operating in turbofan mode.

Determining the variation in vehicle heat loads throughout a mission profile is the first step of the vehicle thermal management evaluation. Next, it is necessary to examine the fuel flow rate changes during the various mission segments to determine if the fuel as a heat sink has the potential to absorb the changing vehicle cooling requirements. The total fuel flow rate for the vehicle varies over the mission duration as illustrated in Figure 9. The shape of the fuel flow rate curve indicates that higher heat sink potential exists during the outbound portion of the mission and this increased heat sink availabilitv coincides with the highest overall vehicle heat loads. The fuel flow curve also indicates possible problem areas of the mission when fuel flow rates drop significantly during descent portions of the mission.

Figure 9. Fuel Flow Variation During Mission Corresponds to Heat Sink Availability.

The TM model includes the heat load characteristics and fuel flow rate changes as a function of time and calculates the predicted outlet fuel temperatures at each of the heat exchangers. Plotted fuel system temperature profiles can then be examined to evaluate the balance between heat loads and available fuel heat sink. If a mismatch occurs between heat sink availability (low

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fuel flow rates) and the magnitude of the vehicle heat load, fuel temperatures can exceed limits imposed by the fuel supply system thermal requirements or the fuel chemical stability limits. A thermal management technique must be employed to counter any imbalance between vehicle heat loads and fuel heat sink capacity.

Recirculation of Fuel as a TM Technicrue

An increase in available heat sink can Z be achieved by increasing the fuel flow rate through the vehicle heat exchangers. The preliminary thermal management technique selected for the hypersonic vehicle

shown in Figure 3 : wherein, up to 1000 considers utilizing a fuel bypass loop as TlUC (minute,)

lb/min of extra fuel can be recirculated through the subsystem, cowl lip, and engine component heat exchangers and is then returned to the fuselage fuel tank. This recirculation of fuel is not allowed to flow through the engine/nozzle hot structures reactor/heat exchanser due to the Droblems associated with retui.ning converted hydrogen back to the fuel tanks.

The thermal model utilizes four selected fuel svstem temDerature checks to control fuel bqpass flo;. Two temperature control points function to start recirculation of bypass fuel. To meet engine component cooling requirements, engine inlet fuel temperatures below 300°F are desired. Therefore, if fuel temperatures exiting the cowl lip heat exchanger exceed 300°F, the bypass flow is activated to keep engine inlet fuel temperatures below this limit. Also, if the fuel temperature exiting the reactor/heat exchanaer is areater than the 1400'F fuel ~ < - - ~

~~~~~~~

chemical stability limit, bypass flow is introduced to maintain fuel temperatures below this limit. The remaining temperature checks are conducted to prohibit the initiation of bypass flow that results in potential problems. If the temperature of the recirculated fuel exceeds 570°F, the potential exists for conversion of MCH to toluene and hydrogen even without the presence of a catalyst. Therefore, recirculation of fuel is not allowed in instances when recirculated fuel temperatures exceed 570'F. Bypass of fuel back to the tanks causes the bulk fuel temperature to increase. Increased tank fuel temperatures call for increased fuel tank pressures to prevent vaporization of fuel. Increased pressures result in thicker, heavier fuel tank walls. Consequently, the tank fuel temperatures are assumed to be limited to 25OOF. Fuel bypass can not occur if tank temperatures are at or above 250°F.

Fission TM Model Results

The fuel system temperature profiles predicted from the HVTM model are shown in Figure 10 for the entire design mission. Each curve represents the fuel system

Figure 10. TM Model Fuel system Temperature Predictions Reveal Imbalances Between Heat Loads and Available Heat Sink of Fuel.

temperature exiting the fuel tank or heat exchangers as defined in the legend. The curves illustrate how the fuel temperature increases as fuel passes through the vehicle heat exchangers as it flows from the fuel tank to the engine combustor. These results include the control of recirculated bypass fuel as previously discussed. The thermal model results presented in Figure 10 point out the following thermal management characteristics and concerns for hypersonic vehicles :

0 The significant engine/nozzle hot structures heat loads during the outbound cruise (Mach 5.5) result in fuel temperatures above or approaching the fuel stability temperature limit of 1400'F for MCH. Even with maximum bypass fuel flow of 1000 lb/min fuel temperatures exiting the reactor/heat exchanger reach 1900'F at the end of the outbound cruise (t=40 min) . Accurate prediction of engine/nozzle cooling requirements for hot structures is essential since this is the dominant heat load during the outbound cruise.

drastically during the idle descent at the end of the outbound cruise (t=40-50 min), the sudden reduction in fuel flow rate occurring from an abrupt throttle setting change causes fuel temperatures to exceed the fuel stability limit.

0 Although vehicle heat loads decrease

0 Recirculation of bypass fuel during the outbound cruise and idle descent makes the bulk fuel temperature increase. From 30 to 50 minutes, the forward tank fuel temperature rises to 250'F and tank pressurization requirements become an issue.

o The fluctuations in fuel temperatures predicted during the return cruise segment of the mission (t=50-112) result from the bypass flow controls.

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o The idle descent at the end of the mission (t>112 min) and the corresponding reduction in fuel flow causes engine inlet fuel temperatures (fuel exiting the subsystem heat exchanger) to exceed 300'F. This temperature is not compatible with the refrigerant requirements used in the ECS cooling system. Therefore, the aircraft subsystems would actually be under cooled, but the potential exists to use ram air for cooling the aircraft sub- systems during this time.

Although bypass of recirculation fuel is utilized as the thermal management technique for the initial studies conducted, the thermal model results indicate fuel temperatures still in excess of fuel stability limits during the outbound Cruise and the idle descent following cruise. This result along with concerns associated with increased fuel tank temperatures (tank pressurization, fuel vaporization, fuel tank strength/weight requirements), the bypass fuel pump characteristics (size, weight, and power requirements), and the possibility of recirculating reacted MCH containing hydrogen influenced the TM model refinements discussed in the next section.

TM Model Refinement

since bypass of recirculated fuel as a thermal management technique has many associated concerns and increases the fuel system complexity, continued TM studies sought to eliminate the need for bypass fuel flow. Balancing the vehicle heat loads entirely with fuel flowing to the engine and determining how this affects vehicle size and performance were the main goals of the subsequent studies. To accomplish this, analysis focused on the problem areas identified during the outbound cruise and idle descent following cruise. Figure 11 reviews the original fuel system temperature predictions during this "problem segment" of the mission when the fuel stability temperature limit is exceeded.

, O l * " O T A Y I rVrL

:;,,. , ,, .-: l?i? "i/ . . . . . . . . . . . . . . . . . . . . . L

P ,600. . . . . . . .

Figure 11. Thermal Management Challenges Identified During Outbound Cruise and Idle Descent.

: &!&&

Fuel temperatures surpassed the fuel stability temperature limit during the outbound cruise of the mission due to the extremely high cooling requirements of the nozzle hot structures. The current vehicle '2 studied is configured with a single expansion ramp nozzle (SERN). Initial nozzle cooling air flow requirements were estimated by scaling the calculated cooling requirement of a two-dimensional, convergent-divergent (2DCD) nozzle based on the increase in nozzle surface area of the SEW. The 2DCD nozzle has two completely enclosed symmetric expansion surfaces: whereas, the upper nozzle ramp of the SERN is exposed to the ambient surroundings. Therefore, this approach does not account for radiation cooling of the upper nozzle ramp surface of the SEW. A thermal analysis conducted by GEAE predicted that SERN cooling requirements decreased 41% when radiation effects were included.

Engine/nozzle hot structures heat loads were re-evaluated accounting for this radiation cooling along the SERN during the outbound cruise and idle descent. Figure 12 presents the revised total vehicle heat loads compared with the original estimated values. The revised total vehicle heat loads are shown to be lower since radiation cooling effects have been included and less cooling air is required to cool nozzle walls.

. . . . . . . . . . . . . . . . . . . . . . , 0 0 0 . . . . . _ * _ * _ _ _ _ > _ . . . . . . . . .

. . . . . . . . . . . . . . . . . i .; . I . . .i . . ~ . . . , / ' " . . ~,,,,.. :. . : . :

. . . . . . .

. . . . . . . . . . . . ,.

. . . . . . . , . . e , . , , ; , , , , . , . -

2 6 . 2 8 . 3 0 . 32. 3 1 . 36, 38, r e . 12. 11. 4 6 , 4 8 . I O . , ,YE Irni""l..)

J

Figure 12. Accounting for Radiation Cooling from S E W Lowers the Total Vehicle Cooling Requirements during the Outbound Cruise.

Throttle Cu tback Schedule Durina Idle Descent

The other problem area identified relates to the abrupt change in fuel flow occurring at the beginning of the idle descent from cruise. The sudden reduction in fuel flow results from an instantaneous change in throttle setting. The available heat sink is thus reduced faster than the vehicle heat loads and this mismatch causes fuel temperatures to spike above the fuel stability temperature limit (1400'F). d

A solution to this idle descent TM imbalance is achieved by developing a throttle cutback schedule; whereby, the throttle is gradually reduced to lower power lever angles (PIA'S). The scheduled changes in throttle setting are determined from a

2400. , , , I I . . . . . . . . 2200. . . . . . . . . . ..> - - - - -

. . . . . parametric study, i n which, heat loads are Dredicted for various PLA settinss and . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5 m o . j . ' ' , . I ,"El 111.1111" ""*'".'Y"C L I Y I I . I..OO r . . . . . .

hescent altitudes. The corresponding fuel $;:::: . . . . . . . . . . . g,,,, . . . . . . . . . .

flow rates are then utilized to compute the fuel temperatures throughout the fuel supply . . . .

: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . system and the maximum~changes in throttle setting are determined. . . . ~ . . . .: .... :.. ..... : ..... <' . . . . . . . . . ; , , . . . . _e.-.____-_____-_-... . . . 2 0 0 . - _,,L:,l,- ;_:,_,,L,:_.:L.III,I,~.~, . . . . . . . . . . . . . : ,.... . . . ............................................ + ..... i ........... i..... . . The influence of throttle cutback . . . . . . . . . .

26 PA. ,e. 31. J.. ,e , 38. ,a, ,*. '1 1 6 . 4 8 5 0 . T i Y S (mi"",.')

0 . schedule on vehicle trajectory is insianificant as illustrated bv Fisure 13. The Gehicle doesn't slow down ai quickly and the idle descent must be initiated sooner ~i~~~~ 1 4 . Revised TM Model Predictions into the cruise leg. The vehicle burns less using New Heat Loads and Throttle fuel during the outbound cruise but more Cutback are Below Fuel Stability during the descent. The end result is a Limits. vehicle with a slightly lower take-off gross weight (TOGW). The cutback schedule does result in a slight increase in the time to station but this remains under the 50 minute requirement imposed on the mission.

W

Figure 13. Throttle Cutback Schedule Has Insignificant Impact on Mission Trajectory.

pevised TM Model Results

Incorporation of the revised vehicle heat loads and the throttle cut back schedule yield the fuel system temperature predictions given in Figure 14. During the outbound cruise, the fuel temperatures remain below the fuel stability limit since radiation cooling of the SERN has been included and the total vehicle heat load has been reduced. The significant reduction in temperature is due to the MCH temperature- enthalpy characteristics. Implementation of the throttle cutback schedule during the descent from cruise helps maintain fuel temperatures below the stability limit with only an instantaneous temperature spike to the fuel stability limit at the very beginning of the descent. It is felt this

L' short duration at the 1400'F temperature is not significant.

The results of this study indicate that, through careful integration of aircraft subsystem and mission performance requirements,-the fuel heat sink capacity of MCH is adequate to meet the cooling requirements of a Mach 5.5 hypersonic vehicle during the specified design mission.

This study illustrates the need and application of detailed transient thermal analysis techniques required to assess thermal management approaches.

The discussion also emphasizes the involvement of the various disciplines (i.e. avionics, flight controls, thermodynamics, propulsion, structures, etc.) needed to develop feasible thermal management system designs.

Most importantly, the results demonstrate the pay-offs that may be achieved by carefully considering the influence vehicle trajectory (Mach vs. Altitude) has on vehicle thermal management and their mutual impact on vehicle Size.

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References:

1. Hagseth, P.E., Gasner, J.A., et.al., "Inlet and Nozzle Concepts for Advanced Airbreathing Propulsion Systems", WRDC-TR- 91-XXXX, Final Report for Period December 1988-December 1990, General Dynamics

Acknowledsements

The authors wish to acknowledae the

corporation, Fort Worth Division, cbntract NO. F33615-88-C-3006, April 1991.

2. Petley, D.H. and Jones, S.C., "Thermal Management for a Mach 5 cruise Aircraft Using Endothermic Fuel", AIM-90-3284, Presented at the AIAA/AHA/ASEE Aircraft Design, Systems and Operation Conference, Dayton, Ohio, September 17-19,1990.

3. Hirvo, D.H., et.al., Aerodynamic/ Structural Heating with Ablation Computer Program P5613 (ASHAB 11). General Dynamics- convair Division, December 1982.

4. SINDA User's Manual, Revision #3, Lockheed Engineering andManagement Services Company, Inc., Contract NAS 9-15800, March 1983.

5. Wildes, L.B. and Eigel, C.R., "The Air Force Endothermic Fuel Program", Fuels Branch, Fuels and Lubrication Division, Wright-Patterson AFB, October 1990.

6. Lipinski, J., et. al., "Design, Fabrication, and Testing of a Boilerplate Endothermic Methvlcvclohexane Fuel Heat Exchanger ReactorSy&em", WRDC-TR-89-2045, June 1989.

7. Manville Aerospace Company, "Min-KB Thermal Insulation for High Temperature Applications in Nuclear Power, Aerospace and Exhaust Systems", Aerospace Insulation Brochure, 1984.

8. Weiting, A.R., NASA Langley Research Center: and Holden, M.S., Calspan-Univ. of Buffalo Research Center, Buffalo, NY: "Experimental Study of Shock Wave Interference Heating On a Cylindrical Leading Edge at Mach 6 and 8": AIAA-87-1511.

9. Hartsel, J.E., et. al., High Speed Propulsion Assessment Program, GE Aircraft Engines for U.S. Air Force, Contract NO. F33615-85-C-2564.

10. Bergholz, R.F. and Hitch, B.D., "Thermal Management Systems for High Mach Airbreathing Propulsion", AIAA-92-0515, Presented at the 30th Aerospace Sciences Meeting & Exhibit, Reno, NV, January 6-9, 1992.

Fuels (WL/POSF) and Aerodynamics & Acrframc (wL/FIMM) Branches of the Wright Laboratory, Wright-Patterson AFB, Ohio for their support of the work presented in this paper. GE Aircraft Engines also provided much appreciated assistance in evaluation of engine and nozzle heat loads. A special thanks goes to William E. Harrison for his guidance and information regarding endothermic fuel characteristics. Technical monitoring of the project was provided by Clay Fujimura, WL/FIMM.

j

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