GB2243912A - Evaluation of static pressure downstream of a normal shock in hypersonic flight - Google Patents
Evaluation of static pressure downstream of a normal shock in hypersonic flight Download PDFInfo
- Publication number
- GB2243912A GB2243912A GB9010458A GB9010458A GB2243912A GB 2243912 A GB2243912 A GB 2243912A GB 9010458 A GB9010458 A GB 9010458A GB 9010458 A GB9010458 A GB 9010458A GB 2243912 A GB2243912 A GB 2243912A
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- pressure
- total air
- air
- value
- static
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/66—Arrangements or adaptations of apparatus or instruments, not otherwise provided for
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
- G01P5/00—Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft
- G01P5/14—Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft by measuring differences of pressure in the fluid
- G01P5/16—Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft by measuring differences of pressure in the fluid using Pitot tubes, e.g. Machmeter
- G01P5/17—Coupling arrangements to the indicating device
- G01P5/175—Coupling arrangements to the indicating device with the determination of Mach number
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- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Remote Sensing (AREA)
- Measuring Fluid Pressure (AREA)
Abstract
A device for installation in an aircraft for determining gas-state parameters for air downstream of a normal shock for hypersonic flight speeds. These are speeds where true-airspeed is in excess of Mach number = 4. The device comprises a total air-temperature sensor 11, a total air-pressure sensor 12, a processor 13, a display 17 and a control-unit 18. The processor includes an input/output unit 14, a central processing unit 15 and an electronic memory 16, storing data relating to properties of air and to parameter relationships dependent upon total air-temperature and total air-pressure. Based upon the sensed parameters and stored data, the processor calculates the static pressure, the value of which enables quantities to be evaluated that are important for piloting and control of the aircraft. These quantities are output to a pilot display 17 and to a control-unit 18. <IMAGE>
Description
EVALUATION OF STATIC PRESSURE DOWNSTREAM OF A NORMAL
SHOCK IN HYPERSONIC FLIGHT
This invention relates to an apparatus and a method for use in the analysis of air data for hypersonic flight. Specifically, the invention relates to apparatus and a method for calculating static pressure downstream of a normal shock.
A value for the static pressure, downstream of the normal shock caused by an aerospace vehicle in hypersonic flight, is required for the evaluation of
Mach number (defined as the quotient in the division of true airspeed by the speed of sound at ambient air temperature, Ta). For the purposes of this discussion hypersonic flight may be regarded as flight at Mach numbers greater than 4. Values for the static pressure may also be used in the evaluation of other airspeed or ambient air quantities, which might include: trueairspeed; calibrated airspeed; ambient air temperature; pressure height; etc. These quantities like Mach number, may be displayed to a pilot. Some, Mach number in particular, may in addition be signalled to control units, such as those controlling variable geometry devices fitted to the propulsion units of the aerospace vehicle.Mach number signals to the control units result in the variable geometry devices being positioned such that the highest, or near highest, practicable operating efficiency may be obtained from these propulsion units.
The input parameter values needed for this procedure are those for an equilibrium air state downstream of a normal shock, (implying the exclusion of other air states):
(i) Total air temperature, Tt2;
(ii) Total air pressure, Pt2; (iii) Static air pressure, p2.
Conventionally, these input parameter values have all been obtained from data registered by transducers receiving signals output by air-data sensors mounted on the aerospace vehicle for which the value of Mach number is sought. The necessary calculation procedures also use data relating to the properties of air as a real gas. Such data may be obtained, for example, from
Data Item 88025 (with Amendment A) published by
Engineering Services Data Unit ESDU International plc., and "Tables of equilibrium thermodynamic properties of air: Volume II, constant pressure" [Brahinsky, H.S. & BR<
Neel C.A.J, published by Arnold Engineering Development
Center [AEDC-TR-69 89 Vol. II].
The conventional method for deriving the static pressure, p2, downstream of a normal shock requires the installation of a pressure sensor mounted on the aerospace vehicle. This sensor senses a local surface pressure on the vehicle and its value, registered by a pressure transducer, must be corrected by some calibration method to give a value of the static pressure, p2. The evaluation of Mach number is very sensitive to the accuracy of this pressure parameter, both in absolute terms and relative to the value obtained for the total pressure, Pt2. However, the calibration method used conventionally for deriving the static pressure has the disadvantage that it cannot be relied on to provide the required accuracy.The present invention seeks to provide an apparatus for determining the static pressure downstream of a normal shock in hypersonic flight, such that the measured value is highly accurate, and without requiring the use of any such pressure sensor.
According to a first aspect of the present invention, there is provided a device for determining gas state parameters for air downstream of a normal shock in hypersonic flight, the device comprising:
a first sensor, producing a first signal dependent on the total air temperature;
a second sensor, producing a second signal dependent on the total air pressure;
an electronic memory, containing stored data relating to the properties of air and to the relationships between parameters dependent on the total air temperature and total air pressure; and
a processor, for calculating the static air pressure, p2, on the basis of the first and second signals and the stored data.
Preferably, the processor calculates a value for the Mach number or other airspeed or ambient air quantities on the basis of the calculated static air pressure.
According to a second aspect of the present invention, there is provided a method of calculating static pressure downstream of a normal shock in hypersonic flight, the method comprising the steps of:
measuring the total air temperature, Tt2;
measuring the total air pressure, Pt2; deriving from the total air temperature and the total air pressure measurement parameters [ZRT/h]t2 and log[pt2/Pref]r where
Z is the compressibility of the air,
R is the gas constant,
h is the specific enthalpy of the air, and
Pref is the standard atmospheric pressure, and
obtaining from the measurement parameters a value for the static pressure, p2, by means of interpolation between known data.
The notation [...it2 is used herein to denote that the parameters inside the brackets are evaluated for conditions which give total pressure and total temperature downstream of a normal shock.
The invention will now be described, by way of example, with reference to the accompanying drawings, in which:
Figure 1 is a block schematic diagram of apparatus in accordance with the invention;
Figure 2 is a graphical representation of the relationship between gas parameters; and
Figure 3 is a graphical representation of the sensitivity of the accuracy of the calculated downstream static pressure, p2, to changes in temperature.
Figure 1 shows an apparatus for calculating the static pressure downstream of a normal shock. A first sensor 11 is mounted on an aerospace vehicle, for detecting the total air temperature Tt2. A second sensor 12 is also mounted on the vehicle, for detecting the total air pressure Pt2. For conditions where dissociation occurs downstream of a shock, the values of P2, Pt2 & Tt2 are those for the eventual equilibrium state. It should also be noted that references to values of pressure and temperature are to absolute values.
Signals from the sensors 11, 12 are supplied to a processor 13, including an input/output unit 14, a central processing unit 15 and a memory 16. The memory 16 is used to store data relating to the properties of air. Outputs from the processor 13 are supplied to a display 17, visible to a pilot of the aerospace vehicle. Outputs may also be supplied to at least one control unit 18, which is used for controlling variable geometry devices positioned on the vehicle to allow the highest practicable operating efficiency to be obtained from the propulsion units.
On the basis of the first and second input signals received from the sensors 11, 12 respectively, the processor acts to calculate a value for the static pressure downstream of the shock. The calculation procedure comprises the following steps:
Step (1) The first and second signals are interpreted as follows:
From values of pt2 and Tt2, obtain values of a third parameter th/RT]t2 and a fourth parameter Zt2.
The value of R must be compatible with the units used for Pt2 and ht2, and may, for example, be 287.22777 J.kg'l.K'1. The essential real gas data, used for deriving the values of the third and fourth parameters, are stored in the memory 16, and are given in standard reference sources such as Data Item 88025 (with
Amendment A) published by Engineering Services Data
Unit ESDU International plc., and "Tables of equilibrium thermodynamic properties of air: Volume II, constant pressure" [Brahinsky, H.S. & Neel C.A.], published by Arnold Engineering Development Center [AEDC-TR-69 89 Vol. II).
Step (2) Using the values of the third and fourth parameters which were calculated in Step (1), values are derived for a fifth parameter [ZRT/h]t2 and a sixth parameter loglo[lo4pt2/pref]. As is conventional, Z is the compressibility of the gas and h is the specific enthalpy. Compatible units must be used for the reference pressure, the value of which may be taken to be 101324.6 N.m~2, as specified by the International
Organisation for Standardisation. Again the value for
Pref is stored in the memory 16.
Step (3) Now, the values derived in Step (2) are used to obtain (P2/Pt2)*, an approximate value for the ratio of the static pressure to the total pressure, presented for Standard ambient air temperature, again as specified by the International Organisation for
Standardisation. In order to achieve this, it is necessary to obtain the relevant data for the required range of applicability of the device and to the required precision. Figure 2 is a graphical representation of the relationships between [ZRT/h]t2, logl0l104pt2/Pref] and (P2/Pt2)*, for some values of the parameters, while Table 1, appended hereto, contains some of the data from Figure 2 to greater precision allowing accurate measurements to be made over a relatively narrow range.The memory 16 is used to store the data contained in Table 1 or in Figure 2.
If required, a value of (P2/Pt2)* can be obtained by interpolation between the stored data.
Step (4) It is known that the value of is sensitive to changes in the actual ambient air temperature, and so assuming the standard air temperature can result in errors. Where Ta is known or where the value of (Ta- Tastd) is known, even approximately, Table 1 or Figure 3 should be used to derive the resulting error in (P2/Pt2) This error value can then be subtracted from the value derived in
Step (3) to obtain final (corrected) values using the expression:
(P2/Pt2) = (P2/Pt2)* + 0.001 x e* x (Ta-Tastd), where e* is the quoted percentage error for every 100C difference between Ta and Tastd. If Ta or (Ta - Tastd) is not known, (P2/Pt2) can be assumed to be (P2/pt2)* Step (5) The required value is now derived as follows:
Using the value of (P2 /Pt2) from Step (3) or (4).
calculate the value of static pressure from:
P2 = (P2 /Pt2) x Pt2- The value for the static air pressure thus obtained can now be used, in a known way, in the calculation of other hypersonic flight parameters, such as the Mach number, the true airspeed, calibrated airspeed, ambient air temperature, pressure height, etc.
EXAMPLE
In this example, the measured values given for the first and second signals are:
Tt2 = 2040 K, Pt2 = 8048.5 Nm'2 , so that log10[Pt2/Pref] = -1.1.
Following the method described earlier the value of P2 is derived as set out below:
Step (1) From the stored data, entering the values
of 2040 K and -1.1 obtain the following values of
third and fourth parameters:
[h/RT]t2 = 4.0026.
Zt2 = 1.00074.
Step (2) [ZRT/h]t2= 0.25002.
loglo[lo4pt2/pref] = 2.9.
Step (3) Stored data corresponding to data from
Table 1 or Figure 2 are used to evaluate the
intermediate result:
(P2/Pt2)* = 0.9108.
Step (5) Using Step (3) The value of the fifth
signal is: p2 = 0.9108 x pt2 = 7330.6 N.m'2 However, a noticeable error may have been introduced if the actual ambient air temperature differs significantly from the standard temperature.
Thus, if the value of (Ta - Tastd) is known to be approximately +400C the final (corrected) value of p2 may be evaluated using step (4):
Step (4). From stored data corresponding to those
in Figure 3:
Error in (P2/Pt2) per 100C = -0.03308 Final (P2/Pt2) = 0.9108 - 0.001 x (-0.0330) x (40) = 0.9108 - 0.00132 = 0.90948.
Step (5) The final value of the fifth signal is
then given by: p2 = 0.90948 x Pt2 = 7319.9 N.m-2.
It is noticeable that the difference in the value of P2 obtained using the optional correction step, where there is a large departure from standard ambient air temperature, amounts to 7319.9 - 7330.6 = -10.7
N/m2.
This change would have a significant effect on the value of Mach number, amounting to about 6.5%, compared with the value obtained using the uncorrected calculation procedure.
There is thus disclosed an apparatus which can accurately determine the static air pressure downstream of the shock, and can thus determine the Mach number, and other flight parameters, without requiring a static air pressure sensor. The ranges of applicability are not defined readily in terms of speed and height, but might be regarded as generally between Mach numbers of 4 and 14 for heights between 33 km (110,000 ft) and -76 Km (250,000 ft), approximately, using the data contained in Figure 2.
TABLE 1 (P2/Pt2)* for Ta # Tastd
Loglot 104pt2/pref [ZRT/h]t2 I 0.245 0.250 0.255 0.260 0.265 0.270 3.5 (0.9150 (0.9120) (0.9087) (0.9047 3.4 (0.9148) 0.9118 0.9084 0.9044 - 3.3 (0.9146) 0.9116 0.9081 0.9041 - 3.2 (0.9143) 0.9114 0.9079 0.9038 - 3.1 (0.9141) 0.9111 0.9077 0.9035 - 3.0 (0.9141) 0.9109 0.9075 0.9031 - 2.9 (0.9140) 0.9108 0.9073 0.9028 - 2.8 (0.9139) 0.9107 0.9071 0.9025 - 2.7 (0.9138) 0.9106 0.9069 0.9022 - 2.6 (0.9137 0.9105 0.9067 0.9019 - 2.5 (0.9136 0.9105 0.9066 0.9017 - 2.4 2.3 2.2 2.1 2.0 e*, ERROR IN (P2/Pt2)* PER +100c DEPARTURE OF Ta ---- -0.0330 -0.0460 -0.0660 FROM Tastd NOTES. (1) Values given in brackets, (0.xxxx), are beyond
the ranges which are valid for this Table, but
are presented for use with numerical
interpolation methods.
(2) Precision of values of (P2/Pt2)* is approximately +0.0005.
Claims (6)
1. A device for determining gas state parameters for air downstream of a normal shock in hypersonic flight, the device comprising:
a first sensor, producing a first signal dependent on the total air temperature;
a second sensor, producing a second signal dependent on the total air pressure;
an electronic memory, containing stored data relating to the properties of air and to the relationships between parameters dependent on the total air temperature and total air pressure; and
a processor, for calculating the static air pressure, p2, on the basis of the first and second signals and the stored data.
2. A device as claimed in claim 1, wherein the processor calculates a value for the Mach number on the basis of the calculated static air pressure.
3. A device as claimed in claim 1, wherein the processor calculates a value for the calibrated airspeed or pressure height on the basis of the calculated static air pressure.
4. A device as claimed in claim 2, further including a display whereby the calculated value of the
Mach number may be displayed.
5. A device as claimed in claim 3, further including a display whereby the calculated value of the calibrated airspeed or pressure height may be displayed.
6. A method of calculating static pressure,
P2, downstream of a normal shock in hypersonic flight, the method comprising the steps of:
measuring the total air temperature, Tt2;
measuring the total air pressure, Pt2; deriving from the total air temperature and the total air pressure measurement parameters [ZRT/h]t2 and log[pt2/Pref]r where
Z is the compressibility of the air,
R is the gas constant,
h is the specific enthalpy of the air, and Pref is the standard atmospheric pressure, and
obtaining from the measurement parameters a value for the static pressure, p2, by means of interpolation between known data.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB9010458A GB2243912B (en) | 1990-05-10 | 1990-05-10 | Evaluation of static pressure downstream of a normal shock in hypersonic flight |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB9010458A GB2243912B (en) | 1990-05-10 | 1990-05-10 | Evaluation of static pressure downstream of a normal shock in hypersonic flight |
Publications (3)
| Publication Number | Publication Date |
|---|---|
| GB9010458D0 GB9010458D0 (en) | 1990-07-04 |
| GB2243912A true GB2243912A (en) | 1991-11-13 |
| GB2243912B GB2243912B (en) | 1994-02-23 |
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ID=10675737
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB9010458A Expired - Fee Related GB2243912B (en) | 1990-05-10 | 1990-05-10 | Evaluation of static pressure downstream of a normal shock in hypersonic flight |
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| Country | Link |
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| GB (1) | GB2243912B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5610845A (en) * | 1994-08-30 | 1997-03-11 | United Technologies Corporation | Multi-parameter air data sensing technique |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB893565A (en) * | 1959-06-03 | 1962-04-11 | Kelvin & Hughes Ltd | Improvements in or relating to computing apparatus |
| GB951441A (en) * | 1959-05-06 | 1964-03-04 | Kollsman Instr Corp | Flight instrument correcting system |
| US4319333A (en) * | 1979-12-19 | 1982-03-09 | Smiths Industries Limited | Mach airspeed indicator |
-
1990
- 1990-05-10 GB GB9010458A patent/GB2243912B/en not_active Expired - Fee Related
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB951441A (en) * | 1959-05-06 | 1964-03-04 | Kollsman Instr Corp | Flight instrument correcting system |
| GB893565A (en) * | 1959-06-03 | 1962-04-11 | Kelvin & Hughes Ltd | Improvements in or relating to computing apparatus |
| US4319333A (en) * | 1979-12-19 | 1982-03-09 | Smiths Industries Limited | Mach airspeed indicator |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5610845A (en) * | 1994-08-30 | 1997-03-11 | United Technologies Corporation | Multi-parameter air data sensing technique |
Also Published As
| Publication number | Publication date |
|---|---|
| GB2243912B (en) | 1994-02-23 |
| GB9010458D0 (en) | 1990-07-04 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| 732E | Amendments to the register in respect of changes of name or changes affecting rights (sect. 32/1977) | ||
| PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 19940523 |